(navigation image)
Home American Libraries | Canadian Libraries | Universal Library | Community Texts | Project Gutenberg | Children's Library | Biodiversity Heritage Library | Additional Collections
Search: Advanced Search
Anonymous User (login or join us)
Upload
See other formats

Full text of "Propulsion subsystem. engineering analysis report, Gemini-Agena target vehicle"

COPY ii.. 



n 



i? 



.',SC-A604141 SP-129-64-17 30 JUNE 1964 



P' 


mi--- 


I i 


i 


o 


1 ', 
i ! 






i •» 




■%,- 


' < 


i 


< 

*- 


i 1 

' i 
i 1 


1 


„, 





PROPOLSie^l SOiSYSIEri 



■M 






5^ 



s4 






i 



*5 
3 



ENBillEERiUS 
GEMir^l-AGEE^I TARGET VE 



T 
CLE 



JUL 2 8 1964 






TO ^ __., 



CEiSSIPICATION CHANCB 





(NASA-CF-96963) PBOPOLSIOS SOBSTSTEM. N79-76535 

ENGINEERING ANALYSIS 8EP0ET, GEHINI-AGENA 
ITABGET VEHICLE (Lockheed Missiles and Space 
Co-) 26 p Onclas 

00/20 11091 



/ 



/ 



/ / 



-./ 



■ ' '/'//fW/j 



/ 




i >-' S J L S S & SPACE COMPANY 

' ' ■ 1 ',.• i v~, i o n of lockheed aircraft corporation 

sunnyvaue:. California 




,,i^J.l«^(*'----'i*-''^--'-^'' 






ia^ eO?¥ HO^ 



// 



. '). . 



In reply refer to: 

I^5SC/A753133 
Orgn. 67-10 

3 F3J. 



COM! 



N Y 



I 

a 



Subject; 
Tot 



Reference ! 



8 June 1965 



Upgrade Action 

NASA Manned Spacecraft Center 
Attn: C. ¥, Matthews 
Gemini Program Office 
Houston, Texas 




OS ■ i' 



•;^--':\^'^ 



CO 



(a) Confidential U^SC report, A60/n)n, "Gemini 
Propulsion System Engineering Analysis 
Report" dated 30 June 196h 



1. Twenty copies of the referenced report were transmitted 
to your office on 16 July 1961; under D^SC transmittal letter A60li093« 

2. Subsequent to the above dispatch, it was determined 
by the Lockheed >!isEiles & Space Company, Space Systems Division 
Classification Office tha t page 6-lg of the report should be upgrade d 
to Confidential, 

3. In accordance with the above, it is requested that 

your office take ths necessary steps to complete this upgrade action. 



m- 





■^m 




;r 


c::::^ 


• m 


1r> 


■2<i 


-O 


r-m 


cap 


mo 


~!r 


<o 



LXKHESD MISSILES & SPACE COMPAKY 



JRLrro 



/OJ R. Lulis 

^cument Control Station-S 




^ 



I 
0- 




ij^ 



4, GROuP D.VI 



S : C N OF I O C >: H fc E D / . r! C R A f T CORPORATION 



CA 



i /' (■■> c 



919 



LMSC-A604141 SP-129-64-17 30 JUNE 1964 



O 

< 

I 

u 
to 

2 



H 



■ i 



■ \ 



I 
I 
I 
I 



PROPOLSiei^ 

ENGillEERING ANALYSIS REPORT 
GEIINI-ABENA TARGET VEHICLE 

(UNCLASSIFIED TITLE) 





Prepared Under Authority of Paragraph 2.2.2 
Contract AF - 04(695)-129 



This doc 
llB revelation of its con. 



fc «U W«*ic « a Bt:<fa l«p»» o i tlw,Uai»£(i ^\a\e% within 



j^jgJ5j^g5[Bt5=l».r«flJW>«irfl»ri««dK|MWi««,:ivju by Tow" 




APPROVED: 



-^V'-^^L,.^ 




^ J. J. KENNELLY, MANAGER 
MSVP ENGINEERING 
APPROVED: 



^-e^^A-e-vi-^uJKA, 



MISSILES & 



V. L. SHOENHAIR, MANAGER 
MEDIUM SPACE VEHICLE PROGRAMS 

SPACE COMPANY 



A GROUP DIVISION OF LOCKHEED AIRCRAFT CORPORATION 

SUNNYVALE. CALIFORNIA 



^^^Tff^™^""Wiiiwiiftiiiiiiiiiiiii I ^_ 



LMSC-A604141 



FOREWORD 



This Propulsion Subsystem Engineering Analysis Report is produced by 
Lockheed Missiles and Space Company for the National Aeronautics and 
Space Agency and Air Force and is prepared under authority of Para- 
graph 2. 2. 2, Contract AF-04(695)- 129. The document describes and presents 
an analysis of the Propulsion Subsystem for the Gemini -Agena Target Vehicle. 



I 



1 

e 



111 



I 



LOCKHEED MISSILES & SPACE COMPANY 



wtBgBCTW>''g^!r ^w . ;Mj w i a a^^ 



yg^ - ' i., j !y. v?mj^m w G r JfS» ^ *" ' ' ' ^fmwif ^^in m^^i in mf f^^ v - ^^mr f'i m^f T ', ■g gpr*'>y-' 



LMSC-A604141 



CONTENTS 



I 
I 
I 
i 



Section Pcige 

FOREWORD iii 

ILLUSTRATIONS vii 
TABLES xi 

ABBREVIATIONS xiii 

1 INTRODUCTION 1-1 
1. 1 Scope 1-1 
1. 2 Vehicle Configuration and Definition of Subsystem 1-2 

1. 3 Mission Requirements 1-4 
1.4 Performance and Environment Specifications 1-6 

2 DESCRIPTION OF GEMINI-AGENA TARGET 

VEHICLE SUBSYSTEM B 2-1 

2. 1 General Description of Subsystem 2-1 
2. 2 Main Engine 2-5 

2, 3 Propellant Feed and Load Systemi 2-52 
2.4 Gas Pressurization System Assembly 2-64 

3 PROPELLANTS AND PROPELLANT LOADINGS 3-1 

3. 1 Rocket Engine Propellants 3-1 

3. 2 Other Mediums 3-4 
3.3 Propellant Loading 3-6 

4 ENGINE OPERATION 4- 1 

4. 1 Ullage Orientation Control 4- 1 
4. 2 Engine Start Sequence (Flight) 4-9 

4.3 Engine Shutdown (Normal Flight) 4-12 

4.4 Rocket Engine Ground Operation 4-14 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



CONTENTS (Cont) 

Section Page 

5 INSTRUMENTATION 5-1 
5. 1 Introduction 5-1 
5. Z Preliminary Instrumentation Requirements 5-2 
5. 3 Flight Instrumentation — Telemetry 5-9 
5. 4 Agena Status Display Panel 5-24 

5. 5 Pressure Switches 5-36 

6 PRIMARY PROPULSION SYSTEM PERFORMANCE 

AND DESIGN CHARACTERISTICS 6-1 

6. 1 Introduction 6-1 
6.2 Engine Requirennents 6-1 
6. 3 Orifice-Fed Pressurization System 6-3 
6.4 Subsystem Performance in Vacuum Flight 6-5 
6. 5 Engine Vacuum Performance 6-9 

6.6 Nominal Performance Data 6-13 

6.7 Exhaust Duct Thrust 6-14 

6.8 Rocket Engine Performance Prediction (Preflight) 6-14 

6.9 Rocket Engine Performance Analysis (Postflight) 6-23 

6. 10 Estimated Engine Altitude Performance 6-37 

7 PYROTECHNICS 7-1 

7. 1 Introduction 7-1 

7.2 Devices Used 7-1 

8 AEROSPACE GROUND EQUIPMENT 

8. 1 Introduction 8-1 
8. 2 Primary Propulsion System Checkout Console 8-1 

8.3 Start Tank Loading Carts 8-12 

8.4 Auxiliary Flushing Unit (Start Tank Flushing) 8-16 
8. 5 Vacuum Start Tank Bake-Out Ov^en 8-17 
8.6 Main System Flushing 8-17 

BIBLIOGRAPHY B-1 



VI 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



ILLUSTRATIONS 

Figure Page 

Gemini-Agena Target Vehicle Configuration 1-3 

Ascent Sequence of Events for Gennini-Agena 

Target Vehicle 1-5 

Gemini-Agena Target Vehicle Inboard Profile 2-2 

Prinnary Propulsion System Fluid Flow Schematic 2-3 

Rocket Engine — Left Side View 2-6 

Rocket Engine — Right Side View 2-6 

Rocket Engine — View Looking Aft 2-7 

Rocket Engine Electrical Schematic Diagram 2-8 

Electrical Gate Overspeed Circuit Block Diagram 2-9 

Mounting of Motional Pickup Transducer 2-11 

Fuel Start Tank 2-15 

Fuel Start Tank Manual Fuel Bleed Valve 2-17 

High-Pressure Nitrogen Charging Valve (Start Tank) 2-17 

Fuel Start Tank Fill and Drain Valve 2- 19 

Oxidizer Start Tank 2-20 

Oxidizer Start Tank Fill and Drain Ciieck Valve 2-21 

Fuel and Oxidizer Venturi and Filter Assembly 2-23 

Oxidizer Dual Check Valve 2-24 

Fuel Dual Check Valve 2-25 

Gas Generator Fuel Solenoid Valve 2-27 

Gas Generator Oxidizer Solenoid Valve 2-29 

Turbopump Assembly 2-33 

Cross Section of Fuel Pump 2-34 

Cross Section of Oxidizer Puinp 2-34 

Vapor Pressure Versus Temperature 

(UDMH and IRFNA) 2-36 



vii 



LOCKHEED MISSILES & SPACE COMPANY 



K 


-1 


1 


-2 


2- 


-1 


2- 


-2 


2- 


-3a 


2- 


-3b 


2- 


■4 


2- 


-5 


2- 


■6 


2- 


■7 


2- 


-8 


2- 


-9 


2- 


-10 


2- 


■11 


2- 


■ 12 


2- 


■13 


2- 


■14 


2- 


15 


2- 


16 


2- 


17 


2- 


18 


2- 


19 


2- 


20 


2- 


21 


2- 


22 



LMSC-A604141 



ILLUSTRATIONS (Cont) 



Figure 

2-23 

2-24 
2-25 
2-26 
2-27 
2-28 
2-29 
2-30 
2-31 
2-32 
2-33 
2-34 
2-35 
2-36 

2-37 
2-38 
2-39 
2-40 

2-41 

2-42 

3-1 
3-2 
3-3 
3 -4a 
3 -4b 
3-4c 
3-4d 
3-4e 



Pilot-Operated Solenoid Valve 
Fuel Valve 

Pilot -Operated Solenoid Valve and Fuel Valve 
Oxidizer Valve 
Thrust Chamber Assenribly 
Thrust Chamber Nozzle Extension 
Cross Section of Engine Injector 
View Showing Engine Injector 
Engine Gimbal Attachment and Movement 
Propellant Tank Assembly- 
Fuel Sump (Propellant Containment System) 
Oxidizer Sump (Propellant Containnaent System) 
Propellant Isolation Valve 

Airborne Portion of Propellant Tank Fill 
and Drain Disconnect 

Propellant Pressurization Systemi 

Airborne Portion of Propellant Tank Vent Disconnect 

Helium Sphere Mounting Method 

Pyrotechnically Operated Helium-Control Valve — 
Normal Condition 

Pyrotechnically Operated Helium-Control Valve — 
First Actuation 

Pyrotechnically Operated Helium-Control Valve — 
Second Actuation 

Specific Gravity Versus Temperature (UDMH and IRFNA) 

Fuel and Oxidizer Start Tank Nitrogen Fill Ports 

Engine Start Tank System 

Fuel to be Loaded at 40° ±5°F 

Fuel to be Loaded at 50° ±5°F 

Fuel to be Loaded at 60° ±5°F 

Fuel to be Loaded at 70° ±5°F 

Fuel to be Loaded at 80° ±5°F 



Page 

2-38 
2-39 
2-41 
2-42 
2-44 
2-46 
2-48 
2-49 
2-51 
2-53 
2-55 
2-56 
2-58 

2-61 
2-65 
2-67 
2-68 

2-70 

2-70 

2-71 

3-2 

3-20 

3-21 

3-24 

3-24 

3-25 

3-25 

3-26 



Vlll 



LOCKHEED MISSILES & SPACE COMPANY 



I 



.aiJ£i^it:^>iL'iSW:; 



17 ,.^i!^reLj«:^^t3:. 



s.i^:. :iiK ssmamisaBf-x- 



LMSC-A604141 



ILLUSTRATIONS (Cont) 



Figure Page 

3-4f Fuel to be Loaded at 90° ±5°F 3-26 

3-4g Fuel to be Loaded at 100° ±5°F 3-27 

3-4h Fuel to be Loaded at 110° ±5°F 3-27 

3r5a Oxidizer to be Loaded at 40° ±5°F 3-28 

3 -5b Oxidizer to be Loaded at 50° ±5°F 3-28 

3-5c Oxidizer to be Loaded at 60° ±5°F 3-29 

3-5d Oxidizer to be Loaded at 70° ±5°F 3-29 

3-5e Oxidizer to be Loaded at 80° ±5°F 3-30 

3-5f Oxidizer to be Loaded at 90 ±5°F 3-30 

3-5g Oxidizer to be Loaded at 100° ±5°F 3-31 

3-5h Oxidizer to be Loaded at 110° ±5°F 3-31 

4-1 Total Propellant Orientation Period Prior to 

Main Engine Ignition Versus Dimensionless 

Vehicle Acceleration 4-3 

4-2 Secondary Propulsion Module —Aft View 4-5 

4-3 Secondary Propulsion System Fluid Flow Schenriatic 4-6 

5-1 Primary Propulsion System Prelaunch Helium Loading 5-4 

5-2 Secondary Propulsion System Prelaunch Nitrogen Loading 5-8 

5-3 Engine Switch Group Nonninal Voltage Levels 5-14 

5-4 Agena Status Display Panel 5-25 

5-5 Primary Propulsion System and Secondary Propulsion 

System Status Display Panel 5-28 

6-1 Fuel Pump Minimum Inlet Pressure Versus 

Propellant Temperature 6-2 

6-2 Oxidizer Pump Minimum Inlet Pressure Versus 

Propellant Temperature 6-2 

6-3 Helium Sphere and Propellant Tank Pressure-Time 

Histories — Squib-Operated Helium-Control Valve 6-4 

6-4 Liquid Head Versus Volume for Propellant Tanks 6-8 

6-5 Variation of Rated Thrust With Altitude " 6-10 

6-6 Estimated Rocket Engine Thrust Limiits Over 

Operating Requirements 6-11 



IX 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



ILLUSTRATIONS (Cont) 



Figure 


6- 


-7 


6- 


-8 


6- 


-9 


6- 


-10 



6-11 

6-12 

6-13 

7-1 

7-2 

7-3 

7-4 

7-5 

7-6 

7-7 
7-8 
8-1 
8-2 
8-3 
8-4 
8-5 
8-6 
8-7 

8-8 

8-9 
8-10 



Variation of Rocket Engine Mininnum Overall 
Specific Innpulse Wich Altitude 

Oxidizer Pump Flow Versus Total Suction Pressure 

Fuel Pump Flow Versus Total Suction Pressure 

Oxidizer Pump Static Pressure Rise Versus 
Flow Rate at Constant Speeds 

Fuel Pump Static Pressure Rise Versus Flow 
Rate at Constant Speeds 

Oxidizer Pump Flow Cavitation Characteristic 

Fuel Pump Flow Cavitation Characteristic 

Typical Shaped Charge — Cone End 

Shaped Charge 

Destructor Initiator 

Shaped Charge and Initiator Assembly 

Retrorocket — Cutaway and End View 

Separation Pin Pusher —Horizon Sensor 
Fairing -Vehicle 

Horizon Sensor Torque Tube Actuating Pin Pusher 

Explosive Shroud Bolt 

Primary Propulsion System Checkout Console 

Oxidizer Start Tank Loading Cart 

Fuel Start Tank Loading Cart 

Start Tank Auxiliary Flushing Unit —(Front View) 

Start Tank Auxiliary Flushing Unit —(Back View) 

Vacuum Bake -Out Oven (Start Tank) 

Pneumatic Portion of Primary Propulsion System 
Checkout Console 

Electric Portion of Primary Propulsion System 
Checkout Console 

Oxidizer Start Tank Servicing Unit 

Fuel Start Tank Servicing Unit 



Page 

6-12 
6-16 
6-17 

6-18 

6-19 

6-26 

6-27 

7-8 

7-9 

7-10 

7-11 

7-15 

7-19 

7-20 

7-22 

8-2 

8-3 

8-4 

8-5 

8-6 

8-7 

8-9 

8-10 
8-13 
8-15 



LOCKHEED MISSILES & SPACE COMPANY 



Table 



LMSC-A604141 



TABLES 



XI 



LOCKHEED MISSILES & SPACE COMPANY 



Page 



2-1 Design Characteristics of Gas Generator Fuel 

Solenoid Valve 2-28 

2-2 Design Characteristics of Gas Generator Oxidizer 

Solenoid Valve 2-30 

2-3 Propellant Isolation Valve Operating Characteristics 2-60 

2-4 Heliunn Sphere Specifications 2-69 

3-1 Cleanliness Requirements of Pneumatic Fluid Media 3-4 

3-2 Transient- and Non-Impulse Propellants Initial Burns 3-7 

3-3 Transient- and Non-Impulse Propellants Final Burn 3-8 

3-4 Residuals (Trapped Non-Impulse Propellants) 3-9 

3-5 Nonninal Tank Volumes 3-9 

3-6 Quantity of UDMH to be Loaded Into Fuel-Start Tank 3-19 

3-7 Quantity of IRFNA to be Loaded Into Oxidizer-Start Tank 3-23 

5-1 Primary Propulsion System Prelaunch Instrumentation 

Requirements (Landline) 5-2 

5-2 Secondary Propulsion System Prelaunch Instrumentation 

Requirements (Landline) 5-7 

5-3 Primary Propulsion System Flight Instrumentation 5-9 

5-4 Secondary Propulsion System Flight Instrumentation 5-18 

5-5 Pressure Switches Design Characteristics 5-37 

6-1 Rocket Engine Performance at Altitude 6-9 

6-2 Estimated Engine Altitude Performance and Overall 

Characteristics 6-37 



LMSC-A604141 



TABLES (Cont) 

Table 

7-1 Gemini Pyrotechnics 

7-2 Cartridge Electrical Characteristics 

7-3 Separation Detonator Electrical Characteristics 

7-4 Self-Destruct System Detonator Electrical Characteristics 

7-5 Performance Characteristics of 0.9 KS-500 Retrorockets 

7-6 Nominal Weights and Dimensions of Retrorockets 

7-7 Booster Retrorocket Electrical Characteristics 

7-8 Separation Bolt Detonating Cartridge Electrical 

Characteristics 



Page 

7-2 
7-3 
7-5 
'7-13 
7-14 
7-16 
7-17 

7-21 



Xll 



LOCKHEED MISSILES & SPACE COMPANY « 

. ' J.WJJI4P..W JIV i Jpj i < iS | i j | j y Mjji;p.Mlij..,j<^Wt' l! i^iyjW*w;'^9ilW ! ^»w^ 



LMSC-A604141 



ABBREVIATIONS 

AGE Aerospace Ground Equipment 

AMR Atlantic Missile Range 

ASP Agena Status Panel (Display) 

ATV Agena Target Vehicle 

BAG Bell Aerosystenas Company 

cch Gubic centimieters per hour 

ops Gycles per second 

EMI Electromagnetic Interferance 

°F Degree Fahrenheit 

f Motional pickup transducer pulse rate output 

gpm Gallons per minute 

GGFSV Gas generator fuel solenoid valve 

GGMPS Gas generator manifold pressure switch 

GGOSV Gas generator oxidizer solenoid valve 

GG Gas generator 

IRFNA Inhibited red fuming nitric acid 

Ibf Pounds force 

Ibnn Pounds mass 

LSB Least significant bit 

MON Mixed oxides of nitrogen 

nm Nautical mile(s) 

Np Nitrogen gas 

N^ Turbine speed 

OFPS Oxidizer feed pressure switch 

POSV Pilot-operated solenoid valve 

POHCV Pyrotechnically operated helium.-control valve 

P Pressure 

P„ Thrust chamber pressure 



Xlll 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



ABBREVIATIONS (Continued) 

psia Pounds per square inch absolute 

psid Poiinds per square inch difference 

psig Pounds per square inch gage 

PIV Propellant isolation valve 

POSV Pilot-operated solenoid valve 

ppna Parts per million 

PPS Primary Propulsion System 

PTVA Propellant Test Vehicle Assembly 

QD Quick Disconnect 

rpm Revolutions per minute 

RMD Reaction Motors Division (Thiokol Chenaical Corporation) 

S/C Spacecraft 

SPS Secondary Propulsion System 

TCPS Thrust chamber pressure switch 

TLM Telemetry 

UDMH Unsynrmnetrical dimethylhydrazine 

"^ Volts, peak-to-peak 

vdc Volts, direct current 

WCG Whittaker Controls and Guidance (Division of Telecomputing 
Corporation 



xiv 



LOCKHEED MISSILES & SPACE COMPANY 



l ^^..^)l l | l|l . l W l ^ l |lgp^^ , ^JyJ^^^^^^■^l ll ^^My^lj [p ^W!^^^ ^ 



LMSC-A604141 



Section 1 
INTRODUCTION 

1.1 SCOPE 

This report presents operational modes and performance capability analysis 
of the NASA Agena Target Vehicle (ATV) utilized in the Gemini program. 
The propulsion subsystem B is defined, mission requirements are set forth, 
and performance and environmental specifications are listed in this intro- 
ductory section. 

• Section 2 concerns the overall Primary Propulsion System (PPS), pre- 
sents a breakdovm of the subsystems, and describes in detail the purpose 
and operation of each component. 

• Section 3 discusses the propellants used and the method of loading these 
propellants. Also, the method for calculating the required quantity of pro- 
pellants to be loaded into the main and start tanks is presented. 

• Section 4 describes general engine operation. Ullage orientation, in- 
cluding the need for and means of implementation and other special require- 
ments, is discussed. 

• Section 5 considers instrumentation requirements affecting the operation 
of the PPS. Three categories are covered as follows: 

1. Prelaunch instrumentation which gives "Launch" or "Hold" 
information 

2. Flight instrumentation which, through telemetry, provides infor- 
mation necessary for evaluating total system and flight performance 



1-1 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



3. Instrumentation on the Agena Target Vehicle Status Display Panel, 
which iraparts information concerning the status of the ATV for 
the final docking phase and for PPS and Secondary Propulsion 
System (SPS) operation by the astronaut subsequent to docking. 

• Section 6 discusses the propulsion subsystem performance capability and 
the parameters governing engine performance and power level in vacuum 
flight. The equations used for computing engine-power level are general in 
that they include all significant parameters known at this time. As continued 
experience is gained, periodic adjustment of these equations may be expected. 

• Section 7 describes and details the pyrotechnic devices together with 
pyrotechnic checkout requirements as used in the Agena Target Vehicle. 

• Section 8 of this report briefly lists and describes the Aerospace Ground 
Equipment (AGE) peculiar to the checkout and servicing of the Model 8247 
rocket engine. 

1. 2 VEHICLE CONFIGURATION AND DEFINITION OF SUBSYSTEM 

The Gemini-Agena Target Vehicle design is an adaptation of the basic 
Agena-D vehicle using the alternate Model 8247 rocket engine and additional 
program-peculiar equipment required for the Gemini mission. This ATV 
is essentially divided into (1) the program-peculiar forward auxiliary 
section, (2) the Agena-D forward and midbody sections, and (3) the program- 
peculiar aft section. The intent of this report is to discuss only the systems 
applicable to subsystem B and its operations. 

Vehicle Configuration 

The auxiliary forward section (Fig. 1-1) consists of the auxiliary equip- 
ment rack, the McDonnell Aircraft Company-furnished docking-adapter 
module, and the clamshell nose shroud. The Agena-D forward section 



1-2 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



1- 




UJ 


d: 


h- 


z 


< 


Tt 






(N 
00 


z 


' 


_J 


LLI 




LU 
Q 


z 


3 


O 


< 


3: 
1,, 


2: 


2 




o 

• r-t 
+J 

n3 
• f-i 

o 
o 



J3 
+-> 
0) 

H 

Pi 

0) 



O 






1-3 



LOCKHEED MISSILES & SPACE COMPANY 



■ "Hl"A: " WB»t l ul.i. J. I ,|UJ i MiJJ)M'»IIW BP J I |!l]^Ml. i WiJ.lllir.aUMlJ-.J!gl»>Vt ' W) ' lvn,-M,-Jt i ll;<Lf ' ^^ 



LMSC-A604141 



houses the main equipment bay, and the midbody contains the main fuel and 
oxidizer tanks which supply propellants through a feed and load system for 
the main engine. The Model 8247 multi- start main engine and the smaller 
Model 8250 maneuvering and ullage orientation engines are located in the 
aft section. 

Orbital length of the ATV is approximately 26 feet. Vehicle weight-on-orbit 
is approximately 7200 lb. This weight includes propellants still remaining 
in the main tanks and available for Model 8247 engine operation after the 
Agena achieves orbit. The Gemini-ATV propulsion system consists of the 
following: 

• Model 8247 rocket engine, also known as XLR-8I-BA-13, and its 
controls, mount, gimbals, and titanium nozzle extension 

• Pyrotechnically operated helium-control valve (POHCV) and 
associated pressurization plumbing 

• Fuel and oxidizer feed and load system, including propellant tanks, 
vents, and fill quick disconnects 

• Propellant isolation valves (PIV's) 

• All associated pyro devices and solid-propellant rockets. 

1.3 MISSION REQUIREMENTS 

The propulsion system is designed to inject the ATV into an orbital path 
(Fig. 1-2) which has an altitude varying between l6l nautical miles (nm) 
and 87 nm when measured at a latitude of 28. 34 deg. A minimum of five 
main-engine burns are available to complete the mission requirements; 
one burn is required for injection into orbit and four subsequent burns are 
available to perform orbital plane and phase changes, as required and 
selected from ground stations or from the spacecraft. All launches are 
presently scheduled to take place from AMR Pad 14. 



1-4 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




■—1 
u 

>> 

<u 

OJO 

l-l 
a 
H 

ni 
C 
(U 



0) 

O 
o 



> 



u 
c 

0) 

<p 
w 

<u 
u 

< 

I 



60 

■ 1-1 



I 



B 



1-5 



1 



LOCKHEED MISSILES & SPACE COMPANY 



WMBll^ac; 5«* Wi 



iw«awc*'gB 3 SK «»WT. -..fw;w 



LMSC-A604141 



Once the ATV is in orbit, the PPS provides the thrust necessary for the 
following maneuvers: 

1. Adjusting the orbits so that they are coplanar and making the 
radius of one apside of the Agena orbit equal to the rendezvous 
radius 

2. Circularizing the Agena at the rendezvous radius and proceeding 
to an inertial location which will be a point lying on the line of 
the apsides of the catch-up orbit 

3. At the initial location of the injection point, transfering the Agena 
to a catch-up orbit for placing the vehicle at the rendezvous point 
when specified. 

1.4 PERFORMANCE AND ENVIRONMENT SPECIFICATIONS 

The latest perfornaance and environment specifications applicable to this 
report are as follows: 

LMSC or LAC: 

1414463 Model Specification, Engine, Rocket, Liquid 

Propellant USAF Model XLR-81-BA-I3 (U) 

1067287 Propellant Umbilical Connectors Remote-Operated 

Quick Disconnect 

1412861 Pneumatic Pressure Vessels 

1414804 Pyro Helium Control Valve 

1415273 Valve Propellant Shutoff, Motor -Operated 

1510458 Pi'opellant Vent Umbilical Quick-Disconnect 

Coupling 

1418B Lines and Fitting, Gas and Liquid, Cleaning of 

1414805 Propulsion System Test Requirements for 

Manufacturing and Final Acceptance Tests 

1416537 Requirement Specification Agena Target Vehicle 

Propulsion System Test Requirement 



1-6 



LOCKHEED MISSILES & SPACE COMPANY 

'* '^^lWW I WIWlSMI > 'i»;l>*'W*,«i!!;'!l-WWi ! l l lWiM^ ' > g ! » BJSW ^■l>■^> [! / '? ^ J,'^ a^fl&i«^ta^^^W^^.^V;■^U ' ,■# .»gf^ S 3a »f^■i g :M?^H^^^^^^ 



I 



LMSC-A604141 



Federal: 



BB-N-411A 
MIL-P~25604C 

MIL-P-7254E 



Nitrogen, Liquid and Gaseous 

Military Specification Propellant, unsymetrical- 
dimethyl -hydrazine (UDMH) 

Military Specification Propellant, Nitric Acid. 



I 



B 
I 
I 
I 
1 



1-7 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Section 2 
DESCRIPTION OF GEMINI-AGENA TARGET VEHICLE SUBSYSTEM B 

2. 1 GENERAL DESCRIPTION OF SUBSYSTEM 

A general arrangement of the NASA Gemini-Agena Target Vehicle (ATV), 
including the location and physical relationships of the various propulsion sub- 
system assemblies, is shown in Fig. 2-1, The entire subsystem is organized 
into the following major assemblies: 

• Main Engine - USAF XLR-81-BA-13 (LMSC-1461969) multi- 
start rocket engine 

• Propellant feed and load system 

• Gas pressurization system assembly 

• Secondary propulsion system 

• Pyrotechnics 

In this section, the main engine, the propellant feed and load system, and the 
gas pressurization system will be discussed in detail so that the engine 
operation (Section 4) may be better understood. Refer to Fig. 2-2 for the 
functional relationship of the various individual components. 



2-1 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




o 

u 

Oh 

xs 
u 

o 



0) 

r— < 

u 

> 

4-" 

0) 
DC 
^1 

a 
H 

c 

00 

< 

I 

■r-H 







I 



2-2 



LOCKHEED MISSILES & SPACE COMPANY 



^•^JSews^BirrrmTOBs^ 



iw» u '■ wJii. i iii i BUB.m* > mx .'ft3vrv iffi.wJ!« iJ' JLW* ' .igW !WtB g T 'iMTO » wv g:^i'T*??? ; 's< ^ 



i r?a tPW7^FWgJS ggfi. ' , ^_':':.^ 



I 
I 



B-136 

OXID. TANK TEMP. 

.15° TO*170"F — 



B-8 

OXID. TANK PRESSURE 

0-60 PSIG - 



HELIUM _ 
FILL Q.D. U 



FUEL VENT O.D. 



r 



n 



B-97 

FUEL TANK TEMP. NO. 2 

0.100°F 



^ 




B-9 

FUEL TANK PRESS. 

0-60 P51G 




0-60 P51G / 

Pad 



ORIFICE 
— r~-2-i — 



FUEL 
t3v 



, , . , SQUIBS I &2 Jl SQUIBS 3&4 




I ULLAGE 75.3 FT'' VOLUME 

CONTROL 
TUBE 

\ PROPELLANTTANK 

„^, ULLAG 
.B-96 TD TUBE-' 
■ FUEL TANK 

TEMP. NO. I 

0-100 'F 



PYRO-OPERATED HELIUM CONTROL 
VALVE 

— PYRO ACTUATED 

= HELIUM PRESSURE ACTUATED 




[J OXID. VENT 0.0. 



P B-7 

HELIUM SUPPLY PRESSURE 
0-4000 PSIG 



B-68 

HELIUM SPHERE TEMP. 

0.120^ 



FILTERS 

1 WILL NOT PASS PARTICLES 
GREATER THAN 0.0 10" DIA. 

2 WILL NOT PASS PARTICLES 
GREATER THAN 0.025" DIA. 

3 WILL NOT PASS PARTI CLES 
GREATER THAN 0.0 30" DIA. 



I 



FUEL FILL AND 
DRAIN Q.D. 



i 

...LiAULICS POWER PACKAGE 



PROPELLANT 

ISOLATION 

VALVE (OXIOIZER)- 

. PSI 6.i 



PROPELLANT DUMP LI NES 



• B-2 OXID. PUMP 
INLET PRESSURE 
0-100 PSIG 



■ /^jjzizxn^nninnnTil 



STANDPIPE 

OXIDIZER 
8.4 FT^ VOLUME 



B-32 0XID. PUMP 
INLET TEMP, 
0° TO 100°F 



OXIDIZER PUMP LIP SEAL 
PRESSURIZATION FROM NITROGEN 
ATTITUDE CONTROL SUPPLY TO 5 PSI 

B-71 OXID. PUMP LIP SEAL 

PRESS. (0-30 PSIA) 

8-35 TURBINE SPEED 
24,980 RPM STEADY STATE 



39.1211^ 



TURBINE MANIFOLD 



B-3 TURBINE MANIFOLD 
PRESS. NO. 1 0-750 PSIG 



[I.U7 

OXID. TANK TEMP. (-Z) 

.!5' TO -170°F 



T*[;T TANK TEMP. 
1.5.53'F . 




OXIDIZER FILL 
AND DRAIN VALVE 



I 



'"^'^''Wff^mxt'^^-m 



•WM«;r!!j|iV ' '»WHWl i WJiitv,J » i; i jj.ttii,MW't!W^ ^ ^^ 



LMSC-A604141 






3 



V/ FLEX 



EC AT 950 PSIA 



? 



lb! 



/ <" y / y.'\ /, 



1/4" FLEX HOSE 



B-82 FUEL VALVE 
ACTUATION PRESS, 
0-1500 PSIG 



Sl. 



r 



OXIDIZER FRANGIBLE DISC 
(RUPTURE- 180 PSD 






OXIDIZER VALVE 




B-US OXID INJECTOR 
PRESSURE 
(0-1000 PSIG) 



THRUST CHAMBER 



506 PSIA 



"X -J 



FUEL VALVE 
1/4" FLEX HOSE 




B-6 COMBUSTION 

/ CHAMBER PRESSURE 
' 0-700 PSIG 



THRUST CHAMBER NOZZLE EXTENSION 



T vr 




B-83 THRUST CHAMBER 
SKIN TEMP. 
-50° TO i800°F 



B-91 COMBUSTION 
CHAMBER PRESSURE 
475-550 PSIG 



B-185 NOZZLE EXT. 
SKIN TEMP. NO. 2 
-200 TO +600 "F 



F =15800 LB (NOM) 



B-184 NOZZLE EXT. 
SKIN TEMP. NO. 1 
-200° TO '800°F 



I 
I 
I 
I 



X 



FUEL BLEED VALVE 



' y yy . 



FUEL START TANK 
□. 



FILL AND 
I VALVE 



B-142 FUEL START 
TANK TEMP. 
-10° TO4300°F 





LEGEND 

FUEL 

OXIDIZER 

FUEL RETURN 

GAS 

ORIFICE AND FILTER 

DRAIN, BLEED OR TEST POINT 


1 1 1 1 1 

H. • y 
o / n 




. U 

^P — 





Fig. 2-2 Primary Propulsion System 
Fluid Flow Schemiatic 

2-3 



LMSC-A604r41 



2. 2 MAIN ENGINE 

The XLR-81-BA-13 (Figs. 2-3, 2-4, and 2-5) is a liquid bi-propellant rocket 
engine with a minimuna capability of five starts and a demonstrated capability . 
of fifteen starts under vacuum conditions. This rocket engine consists of the 
following major components: 

• Thrust-chamber assembly 

• Multiple -restart assembly 

• Turbine-pump assembly 

• Overspeed shutdown electronic-gate and cable assembly 

• Turbine -exhaust duct 

• Propellant manifolds 

• Thrust-chamber nozzle extension 

Fuel used is unsymmetrical dimethylhydrazine (UDMH); oxidizer used is 
inhibited red fuming nitric acid (IRFNA). The propulsion system provides 
the thrust necessary to place the ATV into a selected orbit and to accomplish 
major orbital changes. A minimum of five starts is available for performing 
these maneuvers. 

2. 2. 1 Electronic Gate 

Purpose — The electronic gate (Fig. 2-6) serves a multiple purpose. During 
normal engine operation, this device is a junction and power -distribution box 
to the various engine solenoids. In the event of malfunction, it electronically 
senses turbine overspeed and takes immediate remedial action by removing 
power from the engine solenoid valves, thereby terminating engine operation. 

Normal Engine Operation — The Model 8247 rocket engine commences operation 
any time that a voltage (nominally +28vdc) is applied to either pin A or D and 
a ground is applied to either pin B or E of J6000. When either power or 
ground is removed from these pins, engine shutdown occurs. Note from 
Fig. 2-5 that the start signal when applied at the aforementioned pins, goes (1) 

2-5 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



O® ® 










^>^-- 












1 Gas Generator Oxidizer Solenoid Valve 

2 Gas Generator I'ucl Solenoid Valve 

3 Gimbal Ring 

4 Oxidizer Valve 



5 Thrust Chamber Nozzle Extension 

6 Turbine Exliaust Duct 

7 Fuel Start Tank 

8 Oxidizer Start Tank 



Fio-, 2-3a Rocket Engine - Left Side View 




\ I 



1 Thi-ust Chamber Nozzle Extension 

2 Oxidizer Valve 

3 Thrust Chamber Tie Bars 

4 Turbine Exliaust Duct 



Fig. 2- 3b Rocket Engine -Right Side View 



I 



Z-6 



LOCKHEED MISSILES & SPACE COMPANY 



1 

: i 
{ 



LMSC-A604141 











0) 




m 


































> 




>H ;*. 


































•^ 




Ho 
































0) 


> 


>. 


" 1 




























a> 




nl 


T3 


X> 


c t^ 














u 














.5 




> 


O 

C 

o 


E 

HI 

c/1 


U < 






GJ 






01 
> 


> 














^1 c 




<> 


(> 


to 


?, a 






r* 






rt 








01 








o u 

t53 








<: 


^^-i 




> 


rt 

> 






> 


01 
01 






> 

■a 






■n 


O Q) 




o 
CO 

"3 

3 










rt 
> 


■o 

OJ 
01 

iri 


0) 
> 

> 




.r-l 

rt 
Q 


K 

rt 
H 






> 

01 


01 

o 
X 




o 
c 

u 

N 


2 
o 


B S 


c 
a 


11 


rt 
H 


D 

T5 
C 


c 
rt 






c 
rt 






ual Ch 

ump 

iltcr 


m 
Q 


rt 
O 




c 


01 

c 

GJ 

o 


c 

CD 

o 


3 

c 

GJ 
> 


i > 


rt 


rt 


rt 
m 


■rt 

3 

C 


CO 

J., 
o 


Mm 


N 


0/ 


CI, 

E 

3 


izer D 
izer P 
izer F 


3 


o 




-5 '3 


a 


:^j 


to 


CJ 


Tn S 


01 


CI 


Cj 


Cl 




•D 


X) 


Ti) 


<iJ 


T3 -o -a 




a 

Cl 


X 


sS 


3 


rt 


rt 


■1 


C X 








3 


S 


X 












O 


u, o 


O b. W O fi< U< 


U. 


U<OOOl^U,OOOf^ 


K 


y-* 


CM OO 


■<:}* 


in 


U3 


t^ 


CO CT) 


o 


- 


CM 


CO 


■<^ 


in 


CD 


I^ 


CO 


O^ O »-( 
^ CM CM 


CM 
CM 


CO 
CNI 





0- 











V— .>».,< ■■c^-'^ 




\ W-T).-> ( 



l&rf 






/ Xx"t? I- Ik 




"4H 

<: 

•i-i 
O 

o 

■ H 



0) 

d 

bO 

w 

o 
o 



I 

ti 



2-7 



LOCKHEED MISSILES & SPACE COMPANY 



!W%<y^ i fMJ# ; a}§L»g < »i;wi . Wvti^,»«y.a»S ! |p ! a^ 



' •i^»,^ ' y*pw.jjjiw^|ji i ,fi|| 



LMSC-A604141 




a 

U 

a 



s 

U 
CO 
I— I 
ni 
o 



u 

i-H 

w 

(1) 
d 

bO 
P! 
W 



O 
O 

in 

I 

ro 

■H 



2-8 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



I 
I 
I 
I 




(- 


a: 


Q 


Ul 


X 


o 
o 


u 


d; 


iM 


t- 







ce 






lU 


_i 




U 


< 




D 


z 


n 


O 


o 


=> 


Z 




\f' 


o 


L) 


<. 


2 Q. 


(- 





d: 






LU 






s 




3 


O 




Q. 




t- 


Z 


z 
< 

1- 





I 



z 
o 


q: 










< 


5 




z 


o 


CN 


Ul 


LL 


1- 


5 






D 


z 




Qi 


< 




1- 


fy 




Z 


h- 




— 







I 



2-9 



LOCKHEED MISSILES & SPACE COMPANY 



s 

P 

o 
o 

I— I 

pq 



3 

u 

•1-H 

u 

M 

0) 

^> 
O 

.■0 

O 



o 

o 

f— ( 
H 



vO 

I 

CM 



bC 
• 1-1 



MiWtfwitRffaMnrtife"^*"-^"^.'"- ■ 



LMSC-A504141 



through parallel diodes which provide polarity reversal protection to the 
sensitive electric circuitry of the over speed system, and series -parallel 
wired K- 1 and K-2 relay contacts which distribute the signal to the gas 
generator fuel solenoid valve (GGFSV), the gas generator oxidizer solenoid 
valve (GGOSV), and finally the pilot-operated solenoid valve (POSV). 
Various capacitors, diodes, and resistors are used to suppress any radiated 
electromagnetic interference (EMI) to other electronic components in the ATV 
or spacecraft. 

Engine Reset Technique - Prior to each intended operation of the Model 8247 
engine, a short timed-pulse (approximately 50 fisec long) is sent, via a fixed 
capacitive-resistive type circuit, into pin C of J6000. The purpose of 
sending this pulse is to apply a brief resetting signal to both K- 1 and K-2 
relays. In case of turbine overspeed, this feature allows the resetting of 
the electronic gate whereupon another reset signal to obtain normal engine 
operation may be sent at the discretion of either a ground station or the 
spacecraft. This reset signal is sent and terminated prior to intended 
Model 8247 operation in order to preclude the possibility of having a reset 
signal holding the K- 1 and K-2 contacts in the reset position at the same 
time that the engine is experiencing turbine overspeed and sending a 
K- 1 and K-2 open signal to the same contacts. 

Eng ine Overspeed Operation - The circuitry and operation of the turbine over- 
speed portion of the electronic gate is described by the block diagram 
(Fig. 2-6). 

The turbopump gear -case -mounted, magnetically sensitive motion pickup 
transducer (Fig. 2-7) senses turbine speed by noting the variance influx density 
as the four lightning holes in the oxidizer gear pass the pickup head. These 
holes vary the air gap and change the flux density as they rotate past the sensing 
head, thus causing the transducer to emit four pulses for each revolution of 



2-10 



LOCKHEED MISSILES & SPACE COMPANY 



WI4 i ypM|.^.?fe l !W!|pM^ 



LMSC-A604141 



Q 
< 

-J 



i 




(U 

u 

d 

to 

o 

nJ 

a 
o 



•H 

o 



I 



do 
■i-i 



I 
I 



2-11 



LOCKHEED MISSILES & SPACE COMPANY 

' ';w ii jj,.j» ! )iii.i^i.wiiBt»j«i i w p ^,MJju i La;» ' w ^ ^^ 



LMSC-A604141 



the oxidizer gear. The relationship of turbine speed (N ) in revolutions per 
minute (rpm) to transducer output (f ) in cycles per second (cps) is as 
follows: 

(1) Ratio of turbine wheel to oxidizer gear is 74:43. 

(2) N s (f ^ ) (cycles/sec) (74/43) (60 sec/min) (rev/cycle) 

therefore, 

^t ^ ("-^l^^^^out) 

(3) Thus, given a pulse rate of 1142 cps, 
N^ s (1142) (25. 814) s 29, 480 rpm 

The transducer puts out a signal having an approximate sinusoidal wave form 
(Fig. 2-6), which at a frequency of 1140 cps is nonninally 2. 8 volts peak-to- 
peak (Vpp). This transducer output is fed into pins A and B of P6011 and 
through three isolation transformers (Figs. 2-5 and 2-6). Transformer T, 
is the input transformer for the electronic gate circuitry whereas T^ and T, 
are output transformers to the ATV Status Display Panel and telenaetry, 
respectively. These isolation transformers are used to preclude the pos- 
sibility of a short circuit in vehicle wiring affecting the proper operation 
of the overspeed circuit. 

Overspeed Signal Condition — The Schmidt trigger circuitry picks up the out- 
put signal from transformer T, . When the input signal approaches 1140 
±20 cps (turbine overspeed condition), the voltage level becomes high enough 
to fire the Schmidt trigger. The input signal is then converted into a 
square wave, as shown in Fig. 2-6, and fed into the circuitry of the flip-flop. 
The flip-flop takes the square -wave input, cuts the frequency in half, and 
feeds this signal into the frequency-selective network. 



2-12 



LOCKHEED MISSILES & SPACE COMPANY 



i 



LMSC-A604141 



This frequency-selective network consists of a band pass filter and wave- 
shaping circuitry. The band pass filter allows only a selected-signal frequency 
to pass through, thus cutting out any harmonics and allowing only the basic 
signal to pass. The square wave input is reshaped into a sine wave. 

Then, the signal is fed into a pair of rectifiers and changed into a 2. vdc 
signal which drives the relay drivers in a mode that completes the circuit 
between the 28-vdc vehicle power supply and the "OPEN" relay coils. When 
these relay coils are energized, the relay contacts are tripped, thus breaking 
the circuit to the engine solenoid valves. Removing power from the valves 
allows them to close under the spring load, which terminates the flow of 
propellants to the gas generator and the thrust chamber. 

2. 2. 2 Engine Start System 

Purpose of System -A means of reliably starting the Model 8247 rocket 
engine for initial burn is required. Further restriction is placed upon the 
Model 8247 engine inasmuch as a minimum of four additional engine restarts 
must be available to complete the Gemini-ATV mission, and all of these 
starts must be accomplished under relative zero-g conditions. The previ- 
ously used Model 8096 rocket engine accomplished start and restart by 
means of two solid-propellant charges mounted in containers which were inte- 
gral part of the gas generator assembly. The obvious disadvantage of five or 
more such solid-propellant charges necessitated a new design concept which 
would have reliability, simplicity, and inherent multiple-restart capability. 

Start Tank Operation - Thin-shell engine -mounted tanks were designed for 
containing both liquid propellants and high-pressure gas which are separated 
by a thin moveable-metal bellows. The propellant volume in each tank is 
sufficient to start combustion in the gas generator and maintain this com- 
bustion until the turbopump assembly attains sufficient discharge pressure 
to take over the supply of propellant to the gas generator, as well as 



2-13 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



replenish the propellant removed from the tank to accomplish the starting 
phase. The gas side of the tank is designed to be filled with an inert gas 
(nitrogen) at a sufficiently high pressure to force the propellant through the 
interconnecting plumbing into the gas generator at approximate nominal gas 
generator supply pressures. Once filled, this gas remains trapped and 
reduces in pressure only when propellants are demanded from the start 
tanks during the start phase of engine operation. Then the metal bellows moves, 
forcing liquid from the tank, and thereby increasing the gas ullage space. 
As the turbopump assembly accelerates and produces propellant-discharge 
pressures greater than the pressure then existing on the gas side of the 
start tank, the bellows reverses its direction of movement and the tank 
starts to refill with liquid, while at the same time the gas side pressure rises 
until equilibrium is reached as the gas side pressure approximately equals 
the pump-discharge pressure. The start tank is thusly replenished for 
another restart operation upon later demand. Examination of the engine 
schematic (Fig. 2-2) reveals that upon engine shutdown, the liquid propellants 
are trapped between the dual check valves and the closed gas generator 
solenoid valves. 

Start Tank Material - The tank shells are machined from a type 355 stainless- 
steel closed die forging. Prior to final machining, these tanks are heat 
treated to SCT (sub-zero cooling temperature - 185, 000 to 205, 000 psi 
tensile). 

2. 2. 2. 1 Fuel Start Tank . A description of the engine start system fuel tank 
follows. 

Tank Volumes and Pressures — The fuel start tank (Fig. 2-8) has a total 
nominal volume of 187.9 in. of which 95 in. is the maximum expellable, 
7. 6 in. ^ is trapped residual (i. e. , will not come out with normal draining), 
and 85. 3 in. is gas ullage space. Maximum allowed operating pressure 
is 1875 psig; nominal tank working pressure is approximately 980 psig. 



2-14 



LOCKHEED MISSILES & SPACE COMPANY 



I 



LMSC-A604141 







u 

■t-> 
w 

I— I 

& 



00 



2-15 



LOCKHEED MISSILES & SPACE COMPANY 



^ fc^« l wB■ ^ llUkJ^ l ^g^KM^ ^ l*|] ^!B^g!m!fgg ^^w< | jiWm w ^ '■■ s ' ^ ' g ^^ *«^ ^ *y" 



LMSC-A604141 



Tank design burst pressure is approximately four times the nominal working 
pressure; therefore, personnel can safely work in the immediate area while 
the tank is being pressurized. 

Ta nk Construction - The bellows is a nested ripple-type convolution 
design. When suitably restrained in the longitudinal and lateral directions 
and extended to 10 in. , the bellows withstands a 100-psid internal pressure 
differential; when compressed to 1. 960 in. , it withstands 800-psid external 
pressure differential. The bellows is capable of withstanding a minimum of 
50, 000 cycles (each cycle being an extension and compression as previously 
described). The material used is AM 350 corrosion-resistant steel. Using 
the automatic Tungsten Inert Gas process, the bellows is constructed by 
welding at the convolutions with gas shielding on both sides of the welds. 
Each bellows consists of 65 convolutions. The final bellows assembly is leak 
tight when pressurized with helium to 50 psid internally and extended to 10 in. 
and restrained suitably in the longitudinal and lateral directions. During 
normal tank operation, the pressure differential across the bellows is negligible, 

Auxiliary T ank Valves - Three valves provide access to the start tank. 

They are: 

1. The manually operated fuel bleed valve (Fig. 2-9) which provides 
a means of flushing and functionally checking the start tank. This 
valve has a single, hand-operated, metal-to -metal sealing valve 
stem. The bleed port and the valve actuator are covered by an 
AN-919 pressure cap at all times other than during checkouts or 
loading, thereby providing a redundant pressure seal. 

2. The N^-gas fill valve (Fig. 2-10) which is used to pressurize the 
gas side of the start tank. It is a military standard part, having a 
single poppet with a metal -to -metal seal. An AN-919 pressure 
cap is used as a redundant pressure seal to insure that leakage 
does not occur. 



2-16 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



VALVE SHOWN IN 
OPEN POSITION 




TANK 
"END 



Fig. 2-9 Fuel Start Tank Manual Fuel Bleed Valve 



-RING • PACKING, BACKUP, TEFLON 
MS28782'3 



O RING 
AN6227 



.071 FULL 
POSITION W 
2-1/4 TURNS 
NUT 



VALVE ACTUATING 
NUT 




-CAP 
MS20813-1 



Fig. 2-10 High-Pressure Nitrogen Charging Valve (Start Tank) 



2-17 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



3, The fill and drain check valve (Fig. 2-11) which accomplishes 

initial loading of the start tanks. Fill is accomplished through the 
single-poppet check portion of the valve. The remainder of the fluid 
passages are open, serving as a junction for lines coming from the 
fuel pump and the start tank and going to the gas generator. The fill 
port is capped with an AN-919 pressure cap to provide a redundant 
seal at all times other than during tank loading and checkout. 

2. 2. 2. 2 Oxidizer Start Tank. A description of the engine start system oxi- 
dizer tank follows. 

Tank Volumes a nd Pressures — The oxidizer start tank (Fig. 2-12) has a total 

— — ^ 2 

nominal volume of 31 in. of which 13. 8 in. is the maximum expellable, 
2. 2 in. ^ is trapped residual (i.e., will not come out with normal draining), 
and 15. in. is gas ullage space. Maximum allowable operatmg pressure is 
1770 psig; nominal tank working pressure is approximately 975 psig. Tank 
design burst pressure is approximately four times the nominal working pres- 
sure; therefore, personnel can safely work in the immediate area while the 
tank is being pressurized. 

Tank Construction — The bellows is of nested ripple -type convolution design. 
When suitably restrained in the longitudinal and lateral directions and extended 
to 3 in. , the bellows withstands a 100-psid internal pressure differential 
when compressed to 0. 970 in. , it withstands 800-psid external pressure 
differential. The bellows is capable of withstanding a minimum of 50, 000 
cycles (each cycle being an extension and compression as previously de- 
scribed). The material used is type 347 stainless steel. Using the automatic 
Tungsten Inert Gas process, the bellows is constructed by welding at the 
convolutions with gas shielding on both sides of the welds. Each bellows 
consists of 32 convolutions. The final bellows assembly is leak tight when 
pressurized with helium to 50 psid internally and extended to 3 in. and 
restrained suitably in the longitudinal and lateral directions. 



2-18 



LOCKHEED MISSILES & SPACE COMPANY 



I 



LMSC-A604141 



TO GAS GENERATOR 




FROMTURBOPUMP 
DISCHARGE 



LIQUID 
FILL 




LIQUID 
FILL 



START 
TANK 



Fig. 2-11 Fuel Start Tank Fill and Drain Valve 



2-19 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



• GAS FILL 
PORT 




BLEED 
PORT 



FILL AND 
DRAIN PORT 



Fie. 2-12 Oxidizer Start Tank 



Au xiliary Start Tank Valves - Three valves provide access to the start 
tanks. They are: 

(1) The oxidizer bleed valve which is identical to the fuel bleed valve 
(Fig. Z-9) except for thread sizes. 

(2) The N2_-gas fill valve which is identical to the fuel start-tank 
gas-fill valve (Fig. 2-10). 

(3) The fill and drain check valve (Fig. 2-13) which accomplishes 
initial loading of the start tank. Fill is acconnplished through the 
single-poppet check portion of the valve. The fluid passages are 
open to the start tank and the oxidizer pump. The fill port is 
capped with an AN -9 19 pressure cap to provide a red\indant seal 
at all times other than during tank loading and checkout. 



2-20 



LOCKHEED MISSILES 8c SPACE COMPANY 



LMSC-A604141 




^^^^^ 



^^^^^ 



< 



< 



o 



O 



> 

> 



u 
Q 

a 






■!-> 
•H 

• l-l 

o 



CO 



I 



W) 



t 



2-21 



I 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



2.2.3 Cavitating Venturis 

Purpose — The cavitating Venturis provide primary control of engine thrust 
level and are accurately calibrated to regulate flow of oxidizer and fuel to 
the gas generator. Since the gas generator provides power for driving the 
turbine pumps at the speed required to supply the thrust chamiber at rated 
flow, the Venturis indirectly control the engine thrust level by controlling the 
a-mount of propellant to reach the gas generator. 

Venturi Construction — Figure 2-14 shows a typical venturi cross -section. 
Fuel and oxidizer Venturis are similar except for size. Venturis are built 
of stainless steel for corrosion and erosion resistance and are machined 
internally to an extremely smooth finish for high efficiency. The converging 
section is short and the diverging section is relatively long. Throat dia- 
meters vary but are approximately 0.031-in. in diameter for the oxidizer 
and 0. 106-in. in dianneter for the fuel. They have an integrally mounted 
l65-micron filter. 

2.2.4 Dual Check Valves, Fuel and Oxidizer 

Purpose — The dual-poppet check valves (Fig. 2-2) are an integral part of 
the start system. Prior to each engine burn, the start system is maintained 
in a pressurized condition in preparation for the next engine start. The 
purpose of these check valves is to maintain under pressure the propellant 
trapped between the start tank, gas generator solenoid valves, and the 
turbopump discharge port (bootstrap port). This dual-poppet design was 
selected to provide a redundant seal and thus increase reliability of the system. 

Engine Operation — At the initiation of the engine-start signal, the check 
valves are closed, thus containing the propellants within the start system. 
As fuel pump and oxidizer pump bootstrap pressure increase above the 
decaying (blowdown) fuel and oxidizer start tank pressures, the dual check 
valves open and allow propellant flow to sustain gas generator operation and 
recharge the fuel and oxidizer start tanks. 

2-22 



LOCKHEED MISSILES 6c SPACE COMPANY 



LMSC-A604141 



1 
I 
I 
I 




1—1 

B 

0) 

w 
w 

<^ 

u 

0) 

4-> 
r-< 
• H 

a 

■H 

> 

(U 

•H 

Tl 
■■-( 

(§ 

P! 

ni 

r-H 
(1) 



I— ( 

I 






2-23 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



2.2.4. 1 Fuel Dual Check Valve. A description of the fuel dual check valve 
follows. 

Design — The fuel dual check valve (Fig. 2-15) body and poppets are constructed 
of 2024 T-4 alunciinum alloy. The retaining spring is fabricated of passivated 
type 316 stainless steel and is designed to inapart a spring force such that 
upstream off-seating pressure is 5 psig maximum. Dynam.Ic seals are 
virgin TFE teflon on metal seats. Static seals are butyl rubber 805-70 
compound. The unit is designed to have a working pressure of 1875 psig, 
a proof pressure of 2812 psig, and a burst pressure in excess of 3750 psig. 
Maximum pressure drop through the valve at working pressure is 40 psid. 

Checkout —A checkout port is provided in the body and between the two poppets 
for pre-test checkout. This port is pressure capped during normal engine 
operation. Each poppet has a liquid leakage rate of less than 0. 10 cc in 
24 hours when under normal working pressures. 



FLOW 



INLET- 




OUTLET 



-TEFLON 



CHECKOUT PORT 



Fig. 2-15 Fuel Dual Check Valve 



2-24 



e 



LOCKHEED MISSILES & SPACE COMPANY 



-j .» Bi !,« tj j» TO -^r'?^'»;yg'?y 



LMSC-A60414I 



2.2.4.2 Oxidizer Dual Check Valve . A description of the oxidizer dual 
check valve follows. 

Design -The oxidizer dual check valve (Fig. 2-16) body and poppets are con- 
structed of passivated type 304 stainless steel. The retaining spring is 
fabricated of passivated type 316 stainless steel and is designed to impart a 
spring force such that upstream off-seating pressure is 15 psig maximum. 
The poppet seals are virgin TFE teflon 0-rings which butt against metal seats. 
In addition, the start tank side has a teflon seal between the seat and valve 
body. Static seals are of teflon. The unit is designed to have a working 
pressure of 1770 psig, a proof pressure of 2655 psig, and a burst pressure 
in excess of 3540 psig. Maximum pressure drop through the valve at working 
pressure is 30 psid. 



OUTLET- 




BUTYL RUBBER 



CHECKOUT PORT 



Fig. 2-16 Oxidizer Dual Check Valve 



2-25 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Checkout —A checkout port is provided in the body and between the two 
poppets for pre-test checkout. This port is pressure capped during nornrial 
engine operation. Each poppet has a liquid leakage rate of less than 0. 10 cc 
in 24 hours when under normal working pressures. 

2.2.5 Gas Generator Solenoid Valves 

Purpose — The gas generator solenoid valves are spring-loaded normally closed 
electrically actuated valves, which, when open allow flow of propellants to 
the gas generator. These valves are maintained in an electrically energized 
open condition throughout the engine burn. Engine shutdown is initiated by 
removing power froin the GG solenoid valves, which close under spring and 
dynamic loads and shut off propellant flow to the gas generator. 

2.2.5. 1 Gas Generator Fuel Solenoid Valve . A description of the gas generator 
fuel solenoid valve (GGFSV) follows. Refer to Fig. 2-17. Arrows show 
propellant flow passages; letters indicate the following: 

A. Stainless Steel Liner — encapsulates the fluid passages 
to prevent any internal fuel (UMDH) leakage 

B. Spring-Loaded Poppet Assembly — controls fuel flow to the 
gas generator; also a part of the poppet assembly is a ferrite 
core for electrical actuation 

C. Seat Assembly —retains the teflon seat and spring which 
reduces shock on closure 

D. Teflon Seat — provides the actual seal to shut off the oxidizer 
flow to the gas generator 

E. Actuation Gap —indicates the amount of travel available for 
poppet actuation. 

Actuation occurs when the engine-start signal is applied. The solenoid coils 
are energized, and the poppet retracts the annount at (E). This opens up a 
fluid passage between the seat assembly (C), the seat (D), and the poppet 
assembly (B) - allowing fuel to flow from the start tank into the gas generator. 

2-26 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604I41 




> 
> 

O 

d 
o 

O 

w 

1—1 
(D 

1:1 
tx^ 

>H 
O 
+-> 

Jh 

d 

o 

CO 

o 



■—I 
•i-t 



2-27 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



The valve remains open as long as the engine -start signal is applied. "When 
the start signal is removed, electrical power is removed from the solenoid 
coils — allowing the poppet to close under spring and dynamic load and 
shutting off flow of fuel to the gas generator. The seat assembly (C) contains 
a spring to reduce closing shock on the teflon seat. 

Table 2-1 gives pertinent design characteristics of the GGFSV. 



Table 2-1 

DESIGN CHARACTERISTICS OF GAS GENERATOR 
FUEL SOLENOID VALVE 



1. Pressure 

2. Temperature 

3. Solenoid 

4. Response Time 

5. Duty Cycle 

6. Flow 

7. Leakage 

8. Weight 

9. Cycle life 



Operating 

Proof 

Burst 

Operating 

Storage 

Operating Voltage 

Current 

Resistance 

Opening 

Closing 

Gas or Liquid Flow 

No Flow 

Internal at 1875 psig 



1875 psig 

2812 psig 

3750 psig 

-35°to +160°F 

-65° to +165°F 

18 to 30.5 vdc 

3.15 amp at 30. 5 vdc 

9.7fl at 70°F 

0. 100 sec max. at 18. vdc 

0. 02 5 sec max. at 3 0. 5 vdc 

Unlimited "on" at 30. 5 vdc 

10 min "on" at 30. 5 vdc 

1.37 +0. 07 lb/sec 

0. 10 cc/24 hr max. 

3. 4 lb max. 
6000 cycles 



2. 2. 5. 2 Gas Generator Oxidizer Solenoid Valve . A description of the gas 
generator oxidizer solenoid valve (GGOSV) follows. Refer to, Fig. 2-18. 
Arrows show propellant flow passages; letters indicate the following; 

2-28 



I 

B 



LOCKHEED MISSILES & SPACE COMPANY 



g!» VJ.-jJtiM i J; i i»4Bi!LB^ii.ii ii jji].., i j.iij.||^Uil|^^^ 



LMSC-A604141 



I 




> 

> 

•t-i 
o 
c 

<D 

r— I 
O 

CO 

u 

o 

N 

• »-( 

• r-t 

o 

o 

o 

to 

O 



00 



I 






2-29 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



A. Stainless Steel L.iner — encapulates the fluid passages to prevent 
any internal oxidizer (IRFNA) leakage 

B. Spring -Loaded Poppet Assembly — controls oxidizer flow to the 
gas generator; also a part of the poppet assembly is a ferrite cor'e 
for electrical actuation 

C. Seat Retainer —a part of the poppet assembly, retains the 
Kel-F seat (D) 

D. Kel-F Seat —provides the actual seal to shut off the oxidizer flow, 
has a spring load on it in addition to the main-poppet spring load 

E. Actuation Gap — indicates the amount of travel available for poppet 
actuation (provides fluid opening of 0.022 to 0.024 in.). 

Actuation occurs when the engine -start signal is applied. The coils are 
energized, and the poppet retracts the amount at(E). This opens up a fluid 
passage between the retainer seat(C), the seat(D), and the outlet fitting (F)- 
allowing oxidizer to flow from the start tank into the gas generator. The valve 
remains open as long as the engine -start signal is applied. When the start 
signal is removed, electrical power is removed from the solenoid coils — 
allowing the poppet to close under spring and dynamic load and shutting off 
flow of oxidizer to the gas generator. 



For pertinent design characteristics refer to Table 2-2. 

Table 2-2 

DESIGN CHARACTERISTICS OF GAS GENERATOR 
OXIDIZER SOLENOID VALVE 



1. Pressure 

2. Temperature 

3. Solenoid 



Operating 

Proof 

Burst 

Operating 

Storage 

Operating Voltage 

Current 
Resistance 

2-30 



1770 psig 
2655 psig 
3540 psig 
-35° to +160°F 
-65° to +165°F 
18. to 30. 5 vdc 

1 . 7 amp at 30.5 vdc 
18 to 20o at 70 F 



LOCKHEED MISSILES & SPACE COMPANY 



UiM;-.I..~i».i«w-^ arm. .:--. - -. — ,:. -i^ 



LMSC-A604141 



4. Response Time 



5. Duty Cycle 



6. Flow 

7. Leakage 

8. Weight 

9. Cycle Life 



Table 2-2 (Continued) 

Opening 
Closing 
Gas or Liquid Flow 

No Flow 



Internal at 1770 psig 
with operating fluid 



0. 100 sec max. at 18. vdc 
0. 025 sec max. at 30. 5 vdc 
Unlimited "on" at 30. 5 vdc 

30 min "on" at 30. 5 vdc 

0.205 lb/sec at 30 p si max. 
pressure drop 

0. 10 cc/24 hr 

1.8 lb 
6000 cycles 



w 



8 
8 
I 
I 



2.2.6 Gas Generator 

Combustion products of the gas generator (GG) drive a single-stage turbine 
which operates the propellant pumps by means of a gear drive. Initial 
combustion in the GG is initiated by hypergolic ignition of propellants stored 
in the start tanks. These start tanks supply a sufficient quantity of pro- 
pellants to support combustion in the gas generator until turbopump bootstrap 
takes place and flow from the propellant pumps reaches the GG combustion 
chamber. 

The gas generator combustion chamber is made of stainless steel and is an 
integral part of the turbine inlet-manifold. A highly fuel-rich mixture of 
0. 150 lb of oxidizer per lb of fuel is used to maintain relatively low gas 
temperature to the turbine. Since the combustion temperature is relatively 
low (approximately 1400 °F), regenerative cooling is not necessary for the 
gas generator combustion chamber. Nominal flow is 1.575 lb/sec 
(1.370 lb/sec of fuel and 0.205 lb/sec of oxidizer). The nominal chamber 
pressure is 485 psia. 



2-31 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



2. 2. 7 Turbopump Assembly 

The turbopump assembly (Fig. 2-19) consists of a single-stage inpulse-type 
turbine, a fuel pump and an oxidizer pump that are both gear -coupled to the 
turbine shaft, and a gear housing that serves as the assembly frame for the 
three major components (Figs. 2-20 and 2-21). Drain cavities between the 
primary and secondary seals are vented overboard. The oxidizer pumo is 
sealed from the gear case by a sliding-ring-type primary seal, a 
double -lip secondary seal, and an ambient vented-drain cavity between both 
seals. To provide maximum protection against the possibility of acid leaking 
past the primary seal into the gear case, low-pressure nitrogen - Freon gas 
from the attitude control system gas bottles (2-14 psig) - is applied between 
the primary and secondary lip seals any time that oxidizer is in the engine. 

2. 2. 7. 1 Turbine. A description of the turbine follows. 

Design and Operation - the turbine is a single-stage impulse-type designed to 
operate nominally at 24, 800 rpm when driven by hot gases from a bi-propellant 
gas generator. The turbine inlet-manifold incorporates the gas generator 
chamber, adapting to the gas generator injector. 

Exhaust gases are ducted overboard through the elliptical exhaust duct which 
accounts for approximately 200 lb of engine thrust. Although circular exhaust 
ducts on Model 8096 engines are scheduled for incorporation on Agena D 
serial number 92 and up, the Model 8247 engine will continue to use the 
elliptical exhaust ducts. The reason for changing to the round exhaust duct is 
that unsymmetrical flow patterns in the elliptical exhaust duct sets up 
disturbing vehicle torques which create a 7. 5-lb control-gas penalty in 
correction of these torques. The round duct is also 2 lb lighter. However, 
the circular duct presents interference problems with the ATV program 
peculiar aft rack, and redesign of the aft rack is not considered feasible. 
Since the Gemini-ATV Attitude Control System gas supply is not critical, the 
extra requirement presents no problem. 

2-32 



LOCKHEED MISSILES & SPACE COMPANY 



i 



I 



i 
I 
I 
I 




LMSC-A604I41 



2-33 



LOCKHEED MISSILES & SPACE COMPANY 



l>^ 



tn 

(0 

< 



o 

u 
H 



o 

I 

ti) 



LMSC-.A604M1 



PRIMARY 
FUEL SEAL 




PUMP 
INDUCER 



OVERBOARD 
DRAIN OUTLET 



FUEL 

SECONDARY 
LIP SEAL 



TURBINE 
WHEEL — 



B 



Fig. 2-20 Cross Section of Fuel Pump 



NITROGEN/FREON 
INLET FOR SEAL 

PRIMARY PRESSURIZATION 

OXIDIZER 

SEAL 



OXIDIZER 
SECONDARY 
LIP SEAL 




OXIDIZER 
SECONDARY 
LIP SEAL 



-OXIDIZER 
SECONDARY 
SEAL CAVITY 



B 
B 
B 



Fig. 2-21 Cros.s Section of Oxidizer P 



ump 



2-34 



LOCKHEED MISSILES a SPACE COMPANY 



L,MSC-A604141 



Z.2.7.2 Propellant Pumps . A description of the propellant pumps follows. 

Design and Operation — The centrifugal propellant pumps (Fig. 2-19) 
incorporate circular castings and straight vane inapellers with the flow 
terminating in venturi-type diffusers. These diffusers are designed to 
cavitate, thus giving the pumps the desirable characteristic of flow rate in 
direct proportion to speed, independent of head, within the operating 
range of the pumps. Although this design is not as efficient as pum.ps of 
conventional design, it eliminates the need for complex head suppression 
devices to control propellant ratio. Consequently, increased reliability and 
reduced weight are achieved with only a small reduction in pump efficiency. 

Integral steel impellers and shafts (Fig. 2-19) are used in both fuel and 
oxidizer pumps. Aluminum inducers are pinned to the pump impellers to 
reduce the required net positive suction head for the pumps. 

Turbopump Head Requirements — The minimum, total inlet head required 
above vapor pressure at 60°F is 32.5 feet (11. 15 psi) for the fuel pump and 
11.5 feet (7. 82 psi) for the oxidizer pump during start and operation. Figure 2-22 
shows the vapor pressure variance with temperature for both propellants. 
At 60°F, minimum inlet pressure is 13. 05 psia for fuel and 9. 80 psia for 
oxidizer. The design speed for the fuel pump is 25, 389 rpm and for the 
oxidizer pump is 14,410 rpm. The nominal total flow fronn the punaps is 
15.27 lb/sec for fuel and 39-33 lb/sec for oxidizer. 

2.2.7.3 T urbopump Assennbly Lubrication System . The turbine gear box 
contains 150 cc of lubricating oil in accordance with Specification MIL-L-7808E. 
During the engine operation, bearing and gear lubrication is maintained by an 
oil mist formed by agitation of bulk oil by the rotating parts. 



2-35 



LOCKHEED MISSILES & SPACE COMPANY 



•/r.v'tymr ' yg'-H*:'^wg'«wB » ' ^ ^gwwf'^-?w . »ri wimjj <w a 'W<w^^ 



LMSC-A604141 





\ 












* 




X 














•\ \ 


v\. 


X 

Q 

=3 












) 














< 

en 




^ 














\ 














\ 














\ 


\ 



LU 
Oi 

=) 
H 

UJ 



s 



s 



§ 



R 



lo -^ n 

visd - 3anss3ad aodVA 



d 
Q 

Hi 
n! 

0) 

ft 

B 

<D 

H 

w 
;3 
to 
u 

> 

in 
fi 
to 
to 

(U 

0^ 

u 
o 
a^ 

> 



(M 

I 

tM 

bi) 

• H 



2-36 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



2.2.7.4 Engine Gimbal Accessory Drive . A fuel tangential discharge port 
is provided on the fuel-pump housing to supply power to the positive dis- 
placement motor -pump hydraulic package. This hydraulic package, in turn, 
supplies hydraulic power for engine gimbaling. The fuel flow rate through 
the hydraulic power package is 6. 75 ±0. 5 gpm at a pump-discharge pressure 
of 1060 psia, but will vary directly proportionally to engine pump differential 
pressure. This flow requirement is satisfied with an orifice which is sized 
to have pressure-drop characteristics of 365 ±5 psi using UDMH with a 
specific gravity of 0. 795, 

Since the hydraulic system maintains almost constant actuation pressure and 
flow, there is negligible effect on the main engine system except for the 
consumption of a sn:iall amount of gas generator propellant to support the 
added load. The increase in gas generator flow rate for powering the 
auxiliary turbopump is computed at approximately 0.052 lb/sec. 

2.2.8 Pilot-Operated Solenoid Valve and Fuel Valve 

Pilot-Operated Solenoid Valve (PQSV) Purpose and Operation — This valve 
(Fig. 2-23) is a normally open solenoid valve which is actuated by the engine- 
start signal, simultaneously with the gas generator solenoid valves. The 
valve is mounted integrally on and controls the opening of the fuel valve. 

Fuel Valve Purpose and Operation — The fuel valve (Fig. 2-24) is a normally 
closed spring-loaded poppet-type valve operated by fuel -actuation pressure 
from the POSV, which admits fuel to the thrust chamber. This fuel valve 
contains a frangible disc to prevent UDMH from entering the thrust chamber 
and is ruptured by 525-psig fuel-pump pressure on the initial engine start. 
Maximum fuel valve opening pressure is 365 psi. 

POSV And Fuel Valve Assembly Operation — Prior to engine start, the POSV 
poppet is in the normally closed position, and the POSV main poppet is in its 



2-37 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



SPRING 




PILOT 
POPPET 

ACTUATION 
PORT 



PORT 2 



PORT 1 



SECTION A-A 



Fig. 2-Z3 Pilot-Operated Solenoid Valve 



1 



2-38 



LOCKHEED MISSILES & SPACE COMPANY 



WiJ!!iwpgW'.!^>iil!i,.iM| j| jji{j|j^ ji ' ifJ lt.<I B i fepi!» p! B!ffli«f M'jSiii^^ 



LMSC-A604141 




Fig. 2-24 Fuel Valve 



2-39 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



normally open position (Fig. 2-25). Upon application of the start signal, the 
POSV solenoid coil is energized which opens the POSV poppet (simultaneously 
with the opening of the GG solenoid valves). As the fuel punap discharge rises 
to operating pressure, fuel-actuation pressure flows through port B of the fuel 
valve, through the actuation chamber, out through port A^, and returns to the 
fuel pump inlet. Simultaneously, fviel enters the POSV through port D; since 
port C is blocked by the poppet, the fuel flows into the POSV main-poppet 
actuation cavity. Further fuel pressure buildup causes the POSV main 
poppet to close, thus shutting off flow through the main fuel-valve actuation 
cavity out port A, thereby allowing pressure to build up in the main fuel-valve 
actuation cavity. The fuel valve opens before this pressure reaches 350 to 375 
psig, shuttling the main fuel-valve poppet and conapressing the main actuating 
spring. On the initial start, the fuel valve opens and then the frangible disc 
ruptures — allowing flow into the thrust chamber fuel-manifold when discharge 
pressure rises to 525 psig. Upon all subsequent engine starts, fuel begins to 
enter the fuel manifold as soon as sufficient discharge pressure, 350 to 375 
psig, is available to open the fuel valve. 

The shutdown sequence is just the reverse of the start sequence. When 
electrical power is removed from the POSV, the POSV poppet closes under 
spring load. Actuation pressure entering port D to the POSV naain poppet 
is bled off through port C and the POSV main poppet opens. This allows the 
fuel-valve actuation pressure entering port B to decay through Port A, 
and the fuel valve closes under spring load. 

2.2.9 Oxidizer Valve 

Purpose and Operation — The main oxidizer valve (Fig. 2-26) is a normally 
closed spring-loaded poppet-type valve, which admits oxidizer to the thrust 
chamber. Oxidizer pump-discharge pressure on the upstream valve face 
forces the valve open against spring pressure; opening pressure is 



2-40 



LOCKHEED MISSILES & SPACE COMPANY B 

"'"'.!*'i"i | H?! l i W !' .H] ' , ' .t.'iJ ^i:.' . i. ji_ i i iyM«i i Mi ut!i,ujH]ij i j CTy''-'^?ag« Mi^ 'L'i a ^'B a^»<w' y. y.vj^^ 



LMSC-A604141 




> 

. — r 

> 
t— I 

•a 

> 

. — ; 
> 

■r-* 
O 

OJ 
. — I 

o 

CO 

OJ 

■!-> 
Oj 

fH 

(D 

O 
I 



Ph 



in 

00 

I 

N 

ti 



2-41 



LOCKHEED MISSILES & SPACE COMPANY 



o 

D- 



lU 



LMSC-A604141 




I- t^ 

a. <=" 

9 X 

a- o 



2-42 



LOCKHEED MISSILES & SPACE COMPANY 



> 
u 

Kl 

o 



I 



I 
I 
i 



rfmm^>*^.-i iJ ' k -p;,ii|;iR.Mii ' -. ■ ' HMt i j, i ^ife^ i j^,,j-v.ijwL i ^pyj(j.^ 



-.. m 'i - ipin'.'"-r~" '."7^"" 



- . -^.^jB^iqfifc - M vrr^.w rr 



^.j;.i3l;.££5'w,-i^-^ 



v.ii*^i«A^JttaiaL.!s,^iiij-aKi^^.^E5-^:tf:aas^^ ■■ .,*s.._-;3:ir:;s-i-.:,,iR:^^ , 




LMSC-A604141 



2Z5 to 300 psig. A frangible disc in the downstream side of the valve 
guarantees zero leakage prior to first engine start. The frangible disc will 
rupture at a pressure of approximately 180 psi. On shutdown, decay of the 
upstream turbopump discharge pressure allows the valve to close under 
spring pressure. 

Effects of Valve Spring Strength on Oxidizer Postflow —Studies were made 
to determine the effect of oxidizer valve closing times on oxidizer postflow. 
For a decrease from the nonninal value of 2.2 sec (based on 130-psi cracking 
pressure) to 0. 5 sec, postflow would decrease from 20 lb to about 7 lb. 
Decreasing the valve stroke afforded insignificant change over the range 
studied; however, increasing the spring force showed significant effects on 
postflow. Minor modifications to the valve increased the cracking pressure 
range from 225 to 300 psig/ Preliminary calculations indicate the total 
shutdown impulse to be 2300 (+600, -900) lb-sec. 



2.2. 10 Engine Thrust-Chamber Assembly 

Design and Operation — USAF XLR-81-BA-13 engine thrust chamber (Fig. 2-27) 
is an 80-percent bell-shape with an expansion ratio of 45:1. This thrust- 
chamber assembly performs satisfactorily throughout a gimbaling orientation 
of plus 5 deg in a square pattern from the associated planes, with a gimbal 
acceleration of 30 rads /sec/sec. The nozzle is regeneratively cooled up to 
the point at which the area ratio is 13. 3:1. The remaining portion of the 
nozzle (13.3:1 to 45:1) is a radiation-cooled titanium extension. 

Thrust Chamber Construction — The thrust chamber is fabricated of 606l 
aluminum alloy, in three sections: the combustion chamber, nozzle throat, 
and divergent nozzle section. Each section is machined with integral cooling 



2-43 



LOCKHEED MISSILES & SPACE COMPANY 



■■ii-.'^i.-.-.ks'.trf.is-.i.^iitiilnik-i'UM'oi- 




LMSC-A604141 



t- 


> 


UJ 


_i 


V 


< 


CJ 


> 


< 




a: 


a 


m 


o 


LU 


z 


> 


LU 


_j 


_l 


< 


o 


> 


CO 


q; 


Q 


m 


LU 


N 


1- 




< 


5 


DC 


X 


Hi 


o 


CL 



t-l 

s 

(U 
CO 

w 

< 

(U 

s 

U 
in 



I 

bio 

•rH 



■i!!^?Uj:^-f)i<i 



2-44 



o 

-J z 

uj < y 

3 K 12 
U. U. Q 



LOCKHEED MISSILES & SPACE COMPANY 




LMSC-A604141 



passages drilled through the walls of the combustion chamber and nozzle 
sections. These passages are drilled in the combustion chamber in an axial 
direction around the periphery. The holes in the throat section are skewed 
to form a hyperboloid, and holes in the divergent nozzle are drilled to conform 
to the contoured configuration. After matching, these sections are heli-arc- 
welded into an integral assembly. The absence of common welds separating 
the propellant passages, through which leakage might occur, is a special 
feature of this integral design, which provides a rugged structure; eliminates 
the need for external joints; and furnishes a smooth interior for reducing 
turbulence, heat-transfer rates, and performance losses. The inside surface 
of the combustion chamber is coated with aluminum oxide. The radiation- 
cooled titanium nozzle extension is bolted to the regeneratively cooled section 
of the nozzle at the point of 13. 3: 1 area ratio. The titanium nozzle extension 
is coated on the inside with an aluminum -oxide refractory ("Rokide") and 
finished on the outside with a high-emissivity surface. This extension is 
reinforced with molybdenum bands and stringers as shown in Fig. 2-28. 
Nozzle extension failure due to the thermal shock incurred at each firing 
is the limiting factor controlling the number of restarts of the main engine. 
The nozzle throat area is approximately 17 sq in. 

Thrust-Ch amber Flow Passages - The oxidizer enters the thrust-chamber 
cooling passages through a manifold ring located slightly aft of the nozzle 
throat. Flow is divided and equalized in this ring by baffles. All of the flow 
traverses to the exit end and then flows forward through the full length of the 
nozzle and chamber and into the injector manifolding. 

The coolant oxidizer passes through 72 holes of 0. 172-in. diameter to the 
exit-end manifold of the divergent section and returns through a separate 
set of 72 holes (also 0. 172-in. in diameter) to the plenum chamber (manifold) 
adjacent to the inlet manifold. From the plenum chamber, the oxidizer 
passes through 75 holes of 0. 116-in. diameter to the transfer manifold for 



2-45 




LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




Fig. 2-28 Thrust Chamber Nozzle Extension 



2-46 



LOCKHEED MISSILES & SPACE COMPANY 



a 
I 
I 

8 
8 
8 



8 
8 
8 
8 
8 




LMSC-A604141 



cooling the throat section. Forward of the transfer manifold, 185 coolant 
holes of 0. 125-in. dianaeter lead to the injector manifolding. 

Thrust-Chamber Injector - The injector (Figs. 2-29 and 2-30) is a triplet 
type with two fuel jets impinging on each oxidizer jet and a ring of fuel 
doublets around the perimeter to provide film cooling. The injector body is 
machined from an aluminum -alloy forging and is heli-arc -welded to the 
thrust-chamber assembly. It has a cavity in which the fuel valve fits; this 
allows the fuel valve poppet to be located close to the fuel injector orifices. 
Free volumes are thus held small, and when the fuel valve opens, fuel 
injection starts with minimum delay and with minimum "hammer" as the 
small void fills with propellant. A similar arrangement is not needed for 
the oxidizer, since oxidizer flow is accelerated gradually. 

2.2.11 Engine Mount and Gimbal As s embly 

D esign and Operation - The engine mount consists of a nickel-plated 
tubular -steel frame for attachment at four points. The thrust-chamber 
assembly is attached to the engine mount on a mutually perpendicular, 
two -axis gimbal system providing vehicle pitch and yaw control. The gimbal 
mount system provides ±5 deg of deflection in a square pattern; however, the 
hydraulic actuators limit movement to ±2 deg, 30 min in a square pattern. 

A ring containing four equally spaced bearings encompasses the thrust 
chamber forward of the injector. Pivots for these bearings are provided by 
close tolerance 0. 7500 -in. -diameter pins, two of which are positioned in lugs 
welded to the forward face of the injector. The pins consist of a bolt -head 
and thread at opposite ends and are maintained in proper position with lock 
nuts. Rotation around these two pins provides deflection in one plane. In a 
similar manner, two additional pins attached to the engine -mount structure 
permit deflection in the other plane. Thrust is thus transmitted through each 




LOCKHEED MISSILES & SPACE COMPANY 




LMSC-A604141 






fn 









-M 




o 




0) 




•r-i 




fi 




H 




(D 




fi 




•H 




bC 


» 


C 


o 


H 


U. _l 




O U. 


Vl 


Z 1- 





o z 


s 


53 




u 

0) 


Q O 


CO 


q: 




Ql 


m 




m 




O 




M 




U 




c^ 




(M 




(M 




• 




bO 




•H 




h 



2-48 




I 
I 



B 



LOCKHEED MISSILES & SPACE COMPANY 




LMSC-A604141 




Fig. Z-30 View Showing Engine Injector 

2-49 



"■1iiJj W '.^ ' g ' --W. 'I M--^^-^^ ' ^ ' !'! '■ ' - '■ ! W1 B ^'pw|P,'^WW**v~ ' -^* ? ^ ■' ';^;^■^'j^'^'^*^ i-;'ff<ffw-i*y^^' 



LOCKHEED MISSILES & SPACE COMPANY 



II I w'l.^ii^-dtflf^ft^i' r"^-^" •''^ ''•^^^^-■■'■■r^'^'--' 



ir.i ^iumjMuMii' «. ■> im-^mE %»^^*a 



LMSC-A604141 



pair of thrust-chamber and engine -mount bearings. Flexible tubing for the 
propellants connects the pumps and thrust-chamber assembly. Figure 2-31 
shows gimbal attachment to the engine mount. 



I 

1 



I 



2-50 



LOCKHEED MISS ILES 8c SPACE COMPANY 



LMSC-A604141 



VIEW 



ENGINE MOUNTING RING 



ENGINE MOUNT 



PITCH 

HYDRAULIC 

ACTUATOR 




PITCH 
MOVEMENT 



YAW 

HYDRAULIC 

ACTUATOR 





7;r rf r T ^ ■ ^ ■ ^^v 

rrrr^rT^^•n^T 

^r^^m^-\^^^^ 

rrrrrmmT\"i 

rrrrrrmiTT* 

rrrrrmTTn"< 



YAW 

HYDRAULIC 

ACTUATOR 



I 
I 

B 
I 
I 
I 




ENGINE THRUST 
CHAMBER NOZZLE 
EXTENSION 




ENGINE THRUST CHAMBER ASSEMBLY 



ENGINE 

MOUNT-AND-GIMBAL 

RING 



VIEW 



B 



MAXIMUM 

DEFLECTION PATTERN 

■OF GIMBALLED ENGINE 

IS INDICATED BY 

BROKEN LINE 



Fig. 2-31 Engine Gimbal Attachment and Movement 



2-51 



LOCKHEED MISSILES & SPACE COMPANY 






LMSC-A604141 



2. 3 PROPELLANT FEED AND LOAD SYSTEM 

The fuel and oxidizer feed and load systems consists of the following major 
components: 

• Propellant tank assembly 

• Propellant containment systein 

• Propellant isolation valves (PIV's) 

• Propellant fill and drain quick disconnects 

• Propellant fill bellows 

• Propellant pump-inlet feed bellows 

The feed and load system provides controlled means (1) of loading propel- 
lants into the vehicle tanks prior to launch and (Z) of supplying propellants 
to the engine during flight. 

2. 3. 1 Propellant Tank Assembly 

Purpose — The propellant tank asseinbly (Fig. 2-32) stores liquid propel- 
lants in sufficient quantities to allow accomplishment of total Gemini-ATV 
mission objectives. Tankage consists of a dual-chamber assembly, having 
an integrally mounted passive fuel and oxidizer containment system or 
"sumps. " (See Section 2. 3. 2. ) Fuel and oxidizer chambers are separated 
by a common bulkhead, making necessary the maintenance of a nominal 
2-psid pressure differential of fuel over oxidizer to preclude bulkhead 
reversal. 

Operating Pressures and Nominal Volumes — The nominal operating fuel 

and oxidizer pressure is 55 psig each, v/ith a proof pressure of 6l psig. 

3 3 

Minimum net capacities are 75. 9 ft for the fuel tank and 98.4 ft for the 

oxidizer tank. Maximum weight of the propellant tank assembly alone is 

315 lb. 



2-52 



I /^r-u-utm-i N^rccir Ere a. cda/— tr (~r-ikji d a kiv 



LMSC-A604141 




r-i 

s 



ttJ 
H 

C! 

(Ti 
1—1 
r— 1 

(U 

o 



I 

ti) 



2-53 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Z. 3. 2 Propellant Containment System 

Purpose — The propellant containment system, commonly referred to as 
fuel or oxidizer "sumps, " is shown in its relationship to the propellant 
tanks in Fig. Z-3Z. This containment systenn provides improved propel- 
lant scavenging and passive propellant orientation. 

Operation — The concept involves scavenging propellants at depletion and 
containment of sufficient propellants to start and operate the main engine 
until the bulk propellants in the tanks become oriented over the pump inlets. 
A desire to minimize tank-trapped propellant residuals (non-inipulse pro- 
pellants) and to simplify the systems led to the design. 

Figure Z-33 shows the fuel sump configuration and Fig. Z-34 depicts the 

oxidizer sump design. The fuel sump has a nominal capacity of 0. Z5 ft" 

3 
while the oxidizer sump has a capacity of 0. 56 ft . These sumps provide 

approximately two sec of additional burn time by adding their volumes to 

those of the tanks, and by improved scavenging — therefore reducing the 

quantity of non-impulse propellants. Design pressures are identical to 

those for the propellant tanks. 

The containment feature of the sumps is not considered mandatory to ATV 
raission requirements, because the secondary propulsion system is utilized 
for propellant orientation (discussed in Section 4. 1 of this report). However, 
since minimum program peculiar changes were desired to the Standard 
Agena-D Vehicle, the containment system was left intact. This containment 
system offers no detrimental effects to the ATV nor does it influence 
mission requirements or objectives. 



Z-54 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



I 




0) 

-i-i 
w 

■(-> 



■(J 
O 

u 

ni 
i—i 
I— I 

D 

a 
o 

On 

H 
pi 

w 

1— ( 

5 



CO 



Z-55 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




a 



o 
O 

d 
nj 

i-H 

■—1 
(U 
Ph 

o 
a, 



CO 
1:1 

X 

o 



I 

bb 



2-56 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



2. 3. 3 Propellant Isolation Valves 

Operation - The oxidizer and fuel propellant isolation valves (PIV's) are 
motor -operated blade valves located between the propellant tank outlets 
and the propellant pump inlets (Fig. 2-2). In the open position, these 
valves allow filling, draining, and flushing of the tanks through an integral 
fill port located on the engine side of the blade. When closed, the PIV's 
isolate the turbopump inlets and fill lines from the propellant tanks and 
vent engine-entrapped residual propellants through an integral vent port 
which opens only after the valve is fully closed. The vented propellants 
are routed through lines running parallel to the turbine- exhaust duct and 
terminating in nullifiers, thereby precluding the possibility of an increase 
in vehicle velocity after shutdown of the main engine and subsequent closing 
of the fuel and oxidizer PIV's. 

On application of 28 vdc, the valves actuate from the open to closed position, 
or from closed to open, by means of an integral electric motor which is de- 
energized at the fully open and the fully closed positions. The valves in- 
corporate an external visual-position indicator and an electrical connector 
to provide for remote electrical indication and operation. On command, 
these valves cycle repeatedly from open to closed or closed to open. The 
present sequence of events indicates that the ATV will be launched with the 
PIV's open. The propellant isolation valves will close approximately six 
sec after each engine shutdown and upon command reopen four sec prior to 
each engine start. 

These valves were designed for use on Gemini vehicles but will be used for 
all dual- or multiple-burn missions on future Agena vehicles. 

The integrally mounted electric motor {Fig. 2-35) drives a planetary gear 
train which is connected to a shaft. Rotary motion of the shaft is converted 



2-57 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604I4I 



I 




4) 

> 
o 



O 

to 



d 

n) 
.—I 
i-t 

i) 

cu 
o 

!h 

in 
I 



-t-( 

P4 



I 

e 
I 



1- 


LJJ 


is 


q: 


Q 


O 


< 


0. 


_l 


H 1- 


Z 

lu 


to 


Z < 

UJ UJ 

> u. 


> 







2-58 



LOCKHEED MISSILES & SPACE COMPANY 



I 
I 
I 
I 
I 
I 
I 



LMSC-A604141 



to linear motion by means of the crank fixed on the shaft at one end, with 
the engine end sliding in the horizontal slot in the blade. The crank moves 
through an arc while retained in the blade slot. This action forces the 
blade to move in a vertical direction, and thus slide against the seals re- 
tained in the valve body; the blade's sliding action opens and closes the 
flov/ area. The venting mechanism opens only after the valve blade has 
fully closed off the flow area, and closes before the valve blade opens to 
allow^ propellant flow^. 

While on the launch pad, the valves are in the open or "fill" position which 
is maintained at liftoff and in flight prior to first burn. A redundant mom- 
entary open signal is sent to the valves in preparation for engine start in 
order to assure that the valves are open and ready for engine operation. 

The first PIV units were developed as a dual-source item by Whittaker 
Controls and Guidance (WCG) Division of Telecomputing Corporation, and 
by Reaction Motors Division (RMD) of Thiokol Chemical Corporation. 
Both units, identified by the LMSC P/N 1463144-1, were functionally and 
physically interchangeable with the exception that each vendor had a 
basically different venting mechanism, which resulted in a different external 
location for the vent port. Therefore, the RMD valve required longer vent 
tubes than the WCG valve. This difference led to designating the RMD unit 
as P/N 1463144-3. 

Problems w^ith IRFNA salts building up in the WCG-1 valve during long- 
storage periods brought about a redesign of the gear train, which resulted 
in the lowering of the final drive ratio, increasing the torque on the blade, 
and increasing the valve opening time from two to four sec. This latter 
valve is the P/N 1463144-5. The WCG-1 valve will no longer be used. 

In summary, there are two propellant isolation valves which may be used 
on the Gemini-ATV vehicles. These are the RMD- 1463144-3, or the 
WCG-1463144-5. 



2-59 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Table Z-3 presents the operating characteristics of the propellant isolation 
valves. 



Table 2-3 
PROPELLANT ISOLATION VALVE OPERATING CHARACTERISTICS 

1. Flow Characteristics (with IRFNA at 60°F): 

• Pressure drop: 

Inlet to outlet 200 gpm - 0. 50 psi 

Fill-port to inlet-port 60 gpm -5.0 psi 

• Leakage through blade valve 
pressurized to 60 psig o. 10 cch 



• Vent port flow at 5 psia 

2. Actuation time 

(Extreme to extreme) 

3. Actuation voltage 

4. Service life 

5. Structural: 

Proof pressure 
Burst pressure 



1 gpm 

RMD-3 - 2 sec max. 
WCG-5 - 4 sec max. 

22 to 28 vdc unregulated 

1020 cycles 



150 psig 
200 psig 



2. 3. 4 Propellant Fill and Drain Quick Disconnects 

Purpose — The fuel and oxidizer fill and drain quick disconnects (QD's) 
(Fig. 2-36), •which are identical in design, provide access to the propellant 
tanks for loading or draining propellants under controlled conditions. 
These QD's are connected manually and disconnected remotely. 



2-60 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




u 

0) 

P! 
o 
u 
w 

Q 
d 

• r-l 
Q 



a 

+-> 

t— I 
1—1 

<v 
P. 
o 

o 
o 

o 
o 



I 

bio 



I 
I 



2-61 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Operation - The disconnect assemblies are of double-poppet design, with 
one poppet in the airborne and <-he other in the AGE portion of the assennbly. 
As the halves are mated, a seal is established betv/een them. The seating 
ends of both poppets contact each othei- and stay in contact as the poppets 
are forced open by the engaging action of the disconnect assembly. Locking 
is accomplished by^ajball, detent, and sleeve -type ball-retainer arrange- 
ment The disconnect is accom^liMtiiedb^r a ground-controlled piston 
actuator which, when pressurized with gas, moves the sleeve away from 
the ball, thus allowing it to slip out of its detent (which locked the two 
halves together). Internal pressure then acts tCr-i<>rc^-tlve-two halves apart. 
The two poppets move together in the opposite or closing direction, and 
because essentially no free volume surrounds the poppets, virtually no fluid 
is present at the mating interface, thus spillage is minimized as the dis- 
connects separate. If for any reason the pneumatic actuators fail to sep- 
arate as the vehicle lifts off, a redundant lanyard action for separation is 
accomplished by a pull on external lines. 

The airborne portion of the disconnect assembly is shown in Fig. 2-36. 
An auxiliary port on the side is for an optional "Propellant Dunip Kit, " 
which is not used on Gemini-ATV vehicles and is capped. 

2. 3. 5 Propellant Fill Bellows 

The fill bellows, mounted integrally to the disconnect assembly (Fig. 2-2), 
accommodates slight misalignments and contractions of the attached 
components. Within each fill bellows, a conical strainer is positioned to 
prevent the passage of foreign particles larger than 0. 030 in. in diameter. 

2. 3. 6 Propellant Feed Bellows 

The feed bellows (Fig. 2-2), mounted between the propellant isolation 
valves and the turbopump inlets, serves the following functions: 



2-62 



LOCKHEED MISSILES & SPACE COMPANY 



^f}r:^^^»^/!>fT^.;y!mfi^ir^„t^W!' '^7*=*^->' 



^^tB^^^^''*'?*f^r^'9r^l?.y^ ^'^ "'■"'" - ." ^*:* ?- ' ^; ?? ? ^? gg j;^|f'y -^■^■■^•^•^Ffg^ggf^ 



LMSC-A604141 



1. Accommodates minor misalignments of the attached subassemblies 
Z. Isolates engine vibration 

3. Absorbs thermal expansion and contraction of the attached sub- 
assemblies 

4. Absorbs forward movement of the engine during firing 

5. Absorbs expansion and deflection of propellant tanks and sumps 
when filled 

Optional Bellows - Two sets of bellows for fuel and for oxidizer are avail- 
able. The standard set accommodates misalignment from to 0, 140 in. ; 
the alternate set accommodates misalignment from 0. 140 to 0. 280 in. and 
may be required due to a buildup of allowable manufacturing tolerances. 
The choice of bellows varies fromi vehicle to vehicle. 



2-63 



LOCKHEED MISSILES & SPACE COMPANY 



y ' ^ y ?'!ir«*^?'->- -vij:?T-wwwc3HB&3sra?5P«it?^;r?::??^ '■' . 



j-^-if'E Cj r .T ' j i^ ^ y;-' ' 



■^!f^^^'*^^^pw^*N'v■^3l•JiI*-n*'^*v7■«^l*-•■•'•'■•T?^^'™:"^T"*".'^^ 



■7*><y^; ^)HH rn 'M -^-i.. j^ ' Jr'V!m„j«l/ yi VK WI ^ 9 l2S ;^^f^^'^^^ 



LMSC-A604141 



2.4 GAS PRESSURIZATION SYSTEM ASSEMBLY 



Description — The Gemini-ATV pressurization system (Fig. 2-37) supplies 
and maintains the desired pressures in the propellant tanks throughout flight. 
This pressurization system consists of the following: 



(1 
(2 
(3 
(4 
(5 
(6 
(7 



Helium-fill quick disconnect 
Propellant-vent quick disconnect 
Helium sphere (3653 in. capacity) 
Pyro-operated helium-control valve 
Oxidizer-tank pressure switch 
Fuel-tank pressure switch 
Propellant tank AP pressure switch 



Operation — The operation of the gas pressurization system consists of 
opening the helium control valve (first actuation) at some discrete time 
(nonninally 1. 5 sec) after the start signal is sent and closing the oxidizer 
portion of the valve (second actuation) approximately 318 sec later. Control 
of fuel and oxidizer propellant -tank pressures with tinne is altered (1) by 
two 0. 027 -in. -diameter orifices (built into the valve assembly) which control 
the discharge from the gas -storage sphere to the propellant tanks and 
(2) by control of the propellant ullage volumes and pre-pressurization of 
the propellant tanks prior to launch. 

2.4. 1 Helium-Fill and Propellant-Vent Quick Disconnects 

Helium Fill Quick Disconnect — The airborne half of the helium -fill quick 
disconnect (QD) assembly, when coupled with the ground half, provides access 
to the vehicle -propellant pressurization-control system for loading and 
unloading of high-pressure gas. After separation of the halves and prior to 
first actuation of the helium, control valve, the self-sealing airborne half 
prevents the loss of pressurization gas. 



2-64 



LOCKHEED MISSILES & SPACE COMPANY 



■ ■ry'\rT.--T!:?pr?g:?.T-'t>r'. "•■• ■r': '-'W'^iri^T!'^''' 



~ i'^gii yywyp y^y^s^js.^ '• ■ 7- v ?,'^.'iW'':^^7°^J''^?K!^-*5gP?^?*^'T2 



LMSC-A604141 



i 




I 
I 
I 
I 



> 






























u 






























X 






























o 








1- 

u 

M 1 








UJ 














lU 

> 
_l 
< 

> 
_l 

o 

Di 


u 


H 


1- 
U 
UJ 

z 


z 

z 
o 
o 

9 
u 


u 

UJ 

z 
z 
o 
u 

n 


<x 
m 

N 

o 

X 

O 

UJ 

> 

1 


_J 

UJ 
U. 
UJ 

> 

1 


!^- 

Ul :3 

2 UJ 

UJ ce 


UJ 

to 


UJ 

t/5 

UJ 
Qi 




V 

z 
< 






Z 


UJ 


O 


z 


-) 


u 

nr 


Ul UJ 


Ul 


V 




H 


^ 


ro 


o 
u 

2 


z 

o 
u 


UJ 

z 
z 
o 


o 
u 


a 
z 

< 


< 
> 

z 


< 
> 

z 


a: a: 

UJ lu 


Q. 

z 
< 


z 
< 

1- 


Ul 
N 


Ul 
N 


< 

H 

_l 
Ul 

-) 


Ul 
U 
U. 


3 
-J 




u 


u 
a 

H 
Z 
IJI 




z 


o 


o 

1- 


2 3: 


Ul 
N 
Q 

X 


0: 




X 




X 


a: 



UJ 

I 

lu 2 


U 

a 

1 


a 

1 

(J 
o 


Q 

z 
< 

_j 


< 

a: 

Q 

Q 

Z 


< 

-J 

o 

CO 

1- 


< 

_l 
o 
1/1 

t- 


_l -I 
U) Ul 
X X 

DC d; 


-J 
Ul 

u. 
a: 



\. 

_l 

UJ 

u. 




X 

u 


U- 

X 

u 

1- 


DC 
UJ 

Ul 

< 


X (Y 


1 


1— 


■^ 


ir 




^ 


j/. 


Ul UJ 


Ul 


Ul 


T 


i/> 




'^ hi 




y 








< 


< 


u u 


U 


u 


( 1 


III 


Ill 





3 O 

2o 


u. 

3 


UJ 

> 


Ui 
UJ 
N 


ai 

N 


_l 
U- 


_J 
_J 

UJ 


-J 

_J 

UJ 


3 z> 
a a 


Q 
00 


3 


1- 
3: 






z' 


_l 


Q 


Q 


_J 


Q. 


Q- 


z z 


Z 


Z 


to 


t/> 


10 


1^ 


-1 q: 


-J 


UJ 


Ul 


O 


O 


< < 


< 


< 


Q. 


UJ 


Ul 


(N 


UJ >- 


UJ 


D 


X 


X 


3 


a: 


K 


a: q: 


ac 


OH 


fr 


Di 





X Q. 


X 


U. 


o 


o 


u. 


a 


Q. 


H H 


H 


HO 


Ou 


Q. 





— ' CN 


ro 


■«r 


ir! 


^d 


(^ 


00 


t>." 


d •-■ 


CN 


fO 


•«» 


•o 


-0 


r-C 



a 

0) 
■M 
(0 

>> 
w 

el 
o 

■ H 
+-> 

ni 
tsl 
•i-i 
u 
pi 

10 
(0 

<D 

d 

ni 
1—1 
.— I 

(U 

o. 
o 
u 
0. 



I 



i ..!.. l lljj |j ljit 



2-65 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Propellant Vent Quick Disconnect — The airborne half of the propellant vent 
QD, when nnated with the AGE half, pernnits access to the vehicle propellant 
tanks for pressurizing and venting operations without loss of propellant fumes 
or pressurization gas. After separation of the halves, the self-sealing 
airborne portion prevents the loss of pressurization gas and propellant fumes. 

Construction — The propellant-vent quick disconnect (Fig. 2-38) and the 
high-pressure helium disconnect are similar in design. These units are of 
double-poppet design; one half is in the airborne portion and the other is the 
AGE portion of the assembly. Locking is accomplished by a ball, detent, and 
sleeve-type ball-retainer arrangement as described in Section 2.3.4. The 
QD is nornaally disconnected inamediately at liftoff by a pneunnatic actuator 
pulling the lanyard of the connector ground half. If the pneumatic actuator 
should fail, the lanyard itself, when brought into tension by vehicle liftoff, 
causes the ground half to disconnect. If both fail, a section of the airborne 
coupling shears (Fig, 2-38), which allows release without destroying the 
self-sealing capacity of the airborne coupling. 

2.4.2 Helium Sphere 

The helium sphere supplies high-pressure gas for pressurizing the propellant 
tanks during flight. Due to the extended orbital life requirements of the 
Gemini naission, a prograna peculiar 3653 in. sphere is used instead of the 
standard 1612 in. sphere, thereby adding an additional margin of safety and 
assuring the attainment of naission objectives. This sphere is mounted on 
a pedestal in the center of the forward section, as shown in Fig. 2-39. 
Specifications of the sphere are presented in Table 2-4. 



2-66 



LOCKHEED MISSILES & SPACE COMPANY 



>- 
_l 

CO 



LU 
1/5 



LMSC-A604I41 




u 



o 
o 
tn 

Q 

> 

nJ 
H 
+-> 

a 

nj 
t— I 
■— I 
d) 
O. 

o 
0^ 

o 

d 
o 

•r-l 
■1-1 

U 

O 

Oh 

<]J 
Pi 
^^ 

o 

,13 



00 
I 



3 



2-67 



I 



LOCKHEED MISSILES & SPACE COMPANY 



•'^J 'l WIWl.JW. ' JtuWiy ' J.'WA i j i UU|W,M ' B';m!B,. I B\t'/l » L« 









k.-'-'ifi^' 



.y 



^ 6w .7 



■?'" ,^j? ' <» 



r 






/ ■ ' 



\Xi L=fi;i3^ 



-fciSfJSiiliSfc'-Wi'-i^fc™" 



LMSC-A604I41 



E \ T- 




(^ 









I 



I r 






.1 .^ n f '-i 




1 .»■ tl 



■iJ -J 



i 



. / 



-.11 |! 



V 



.1 



U"^: ^ 



K^-' 



k 



..,%^. i< 



*'*ttTi ^. ,>*!, 






NJ 
■^ 






I t 






-? P .if.. 



'"««... } 



i • 






% Y 



#■ 



-sWSfe"*!-.*.-.;..,,^;.? I -~ '•"- 






r: 






■•-•*■*».,,_ 



■ ^( •'. ''^■''**'=a*.-iii'i..,-.<A.tel 



=sA»!* 




Fig. 2-39 Helium Sphere Mounting Method 



I 



2-68 



I 



LOCKHEED MISSILES & SPACE COMPANY 



I 



WV. <^ -,*fftV«L >TO y * sg i w r'y''.'ai?^ ^ f i^. ^ mtt f irw!. } ^'^TT". •■ ^ >*'«??'?J''ff''*Br7W?. « ? ^ y t VtifW-' ! %*W!:'^'Jg; ''*1y g P'^?g'j ffi|g| 



LMSC-A604141 



Table 2-4 
HELIUM SPHERE SPECIFICATIONS 

y^^^^^L 3653 In. ^ 

Working Pressure 2500 psig 

Proof Pressure 3000 psig 

Burst Pressure 4000 pgig 
Mass of Helium 

(loaded at 120°F 

and 2500 psig) 3.144 lb 

2. 4. 3 Pyrotechnically Operated Helium-Control Valve 

P^ifpose-Propellant-tank pressure-time characteristics are controlled by 
the Pyrotechnically Operated Helium-Control Valve (POHCV). High- 



pressure 



gas for the initial pressurization of the propellant tanks is introduced through 
the vent couplings and passages of the POHCV prior to liftoff. During 
flxght, actuation of the valve controls the introduction of pressurization gas 
from the helium sphere into the propellant tanks. 

Figure 2-40 shows the normal or liftoff condition of the POHCV In this 
mode, the helium sphere is filled through the helium-fill disconnect, and the 
propellant tanks are pressurized by means of the vent disconnects. At liftoff 
the vent disconnects and the helium-fill disconnects self seal, and this mode ' 
xs mamtained until Agena first burn. Pressures in the helium sphere and 
propellant tanks are maintained entirely by the disconnect assemblies. 

E1I11AS1}^^^-At 1.5 sec (nominal) after the engine-start signal, squibs 
1 and 2 are fired (Fig. 2-41). High-pressure gas forces the piston to move 
to the left, thus isolating the helium-fill disconnect and allowing high-pressure 

'cluTa'^tl'T T ''' '^"'''^^ "'^^' ^^^^^^^^ ^^°^^^^ P^^^- (^^— -^-at- 
xcally as two). Actuation of the second piston isolates the vent coupling 



2-69 



LOCKHEED MISSILES & SPACE COMPANY 

mmm'^-^ ^ . mmhWJlMJaumuLm ^mmn «, « ^ J. — rira ii.aiy. nM,., 



LMSC-A604141 



HELIUM FILL 
DISCONNECT 



FUEL VENT 
DISCONNECT 



FROM HELIUM 
SPHERE 



i 



3 •»• TO FUEL TANK 



I . I 



LTER —I 



R^ 



%^ 



SQUIBS 
S, 2 






n 



SQUIBS 
3 & 4 



nl 



3 — •» TOOXIDIZER TANK 



OXIDIZER VENT 
DISCONNECT 



Fig. 2-40 Pyrotechnically Operated Helium-Control Valve —Normal Condition 



HELIUM FILL 
COUPLING 



FUEL VENT 
COUPLmO 



FROM HELIUM 
SPHERE 



JPU 






J^ 



RIFICES 



LTER J 



FILTER 




3 — *- TO FUEL TANK 



3 — »- TO OXIDIZER TANK 



OXIDIZER 
VENT COUPLING 



Fig. Z-4I Pyrotechnically Operated Helium-Control Valve — First Actuation 

2-70 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



disconnects and opens up the gas passages from the sphere to the propellant 
tanks. At the same time, helium is applied across a pair of 0. 027-in. 
diameter orifices which control the pressurization-gas flow to the propellant 
tanks. Figure 2-42 presents a typical pressure-time history for the sphere 
and propellant tanks. 

Second Actuation -Approximately 318 sec after the engine-start signal, 
squibs 3 and 4 are fired, thus shutting off the pressurization-gas flow to the 
oxidizer tank (Fig. 2-42). The 318 sec is an optimum time which was 
established from extensive studies that considered the required pump-inlet 
pressures and the necessity of maintaining a 2.0 ±0.2-psid fuel-over-oxidizer 
tank pressure for the astronauts' safety. 



HELIUM FILL 
■ COUPLING 



FROM HELIUM 
SPHERE 




•TO FUEL TANK 



T00XIDI2ERTANK 



OXIDIZER 
VENT COUPLING 



Fig. 2-42 Pyrotechnically Operated Helium-Control Valve - Seco 

2-71 



nd Actuation 



LOCKHEED MISSILES & SPACE COMPANY 



m f't'>^' Hh' f fm>*gT^tiv' -r»pr^ 



LMSC-A604141 



Construction — Interport closure is accomplished with pressure (squib) 
actuated tapered pins which provide a leak-proof metal-to -metal seal. Ports 
are provided so that transducers may be mounted to monitor helium-sphere, 
fuel tank, and oxidizer tank pressures. A 10-micron nominal (25-micron 
absolute) filter is located between the pressurization (helium tank) port and 
orifices, downstream from the helium-fill coupling. 

2.4.4 Tank-Mounted Pressure Switches 

Fuel Tank Pressure Switch — The fuel tank pressure switch is mounted in the 
pressurization system and senses fuel tank-top pressure. The switch is set 
to actuate at 15 ±2 psia. This switch, although part of the pressurization 
system, does not in any way control the pressurization system, and only serves 
to relate information for the A TV -mounted Status Display Panel. A full 
discussion of this panel is included in Section 5 of this report. 

Oxidizer Tank Pressure Switch — The oxidizer tank pressure switch is 
mounted in the pressurization system and senses oxidizer tank-top pressure. 
The switch is set to actuate at 15 ±2 psia. This switch is similar to the fuel 
tank pressure switch and will be discussed in Section 5. 

Propellant Tank AP Pressure Switch - The propellant tank AP switch is used 
to sense the relationship or difference in pressure between the fuel and 
oxidizer tanks. This switch is designed with separate dual-sensing chambers 
which are common only electrically. Like both of the previously mentioned 
pressure switches, this switch is used to supply information to the Status 
Display Panel and will be discussed in Section 5. 



2-72 



T/"Hff > '!i ' ^l|i,H lil l.i i .J i -f ! 



V anmnff^wfufjicimv 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Section 3 
PROPELLANTS AND PROPELLANT LOADINGS 

3. 1 ROCKET ENGINE PROPELLANTS 
3. 1. 1 Fuel Propellant 

Fuel utilized in the Gemini-Agena Target Vehicle (ATV) conforms to specifica- 
tion MIL-D-25604B- 1 unsymmetrical dinnethylhydrazine (UDMH). The fuel 
supplied to the fuel pump inlet may not contain particles greater than 0.030-in. 
in diameter; a filter located on the fuel pump prevents passage of particles 
greater than 0.025-in. in diameter to the gas generator plumbing. 

Safety Considerations . Although UDMH vapor inhalation is the most dangerous 
hazard source, the fuel also can be absorbed into the body through contact with 
the skin and by ingestion. The effects of UDMH are severe irritation of mucous 
membranes of eyes, respiratory passages, and lungs; stimulation of the central 
nervous system resulting in convulsions; and irritation of the gastro-intestinal 
tract; nausea; and emesis. In sufficient quantities, the fuel can cause direct 
destruction of red blood cells. 

Physical Properties. UDMH has a flash point of 34 °F, is highly flammable, 
and is hypergolic in contact with IRFNA. Auto -ignition of the fuel occurs at 
approximately 482 °F. UDMH liquid is subject to oxidation which causes the 
color to change from a clear to yellow. The fuel readily combines with 
moisture in the atmosphere which has a detrimental effect on its use as a 
rocket engine propellant. Spillage of the fuel may be neutralized with water 
and/or dilute sulfuric acid. Specific gravity of UDMH is shown in Fig. 3-1. 



3-1 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 





















/ 


/ 






















/ 


/ 










* 












/ 
























/ 


/ 


[ 


















X 

a 


/ 


/ 




















/ 


/ 


/ 


< 

LL 


















/ 


/ 


/ 


/ 


















/ 


/ 


/ 


/ 



















o 



a: 

H- 
< 

LU 
Q. 

2 



Pi 

t-l 

-d 

PI 

% 
Q 

M 

•a 
+-> 

g 

H 
m 

CO 
M 

> 

•l-J 

> 

u 
O 



O 

a, 



I 

CO 



• t-t 



•0 'dS VNjyi 



00 



•0 "ds Hwan 



3-2 



LOCKHEED MISSILES & SPACE COMPANY 



^^HiyH^tg^ ^;i»M p ^." Mi^-t"T7^pyT"'-'" ' JMt^VS'^ 



K;~Ty "•^7^rgSigg5?gi ' ' ■ 



■r-rCTa'^i?'?i!Pr.fy<«'i»gBtBag«-?^".yCT^i°r'y^^ 



LMSC-A604141 



3. 1. 2 Oxidizer Propellant 

Oxidizer used in the Gemini-ATV propulsion systenn conforms to Specification 
MIL-P-7254E-1, Type IIIB inhibited red fuming nitric acid (IRFNA). Total ' 
solids content of the oxidizer is less than 0.05 percent by weight. The IRFNA 
supplied to the oxidizer pump inlet inay not contain particles greater than 
0.40-in. in diameter; a filter located on the oxidizer pump prevents passage 
of particles greater than 0.0 10 -in. in diameter to the gas generator plumibing. 

Safety Considerations. IRFNA (both liquid and vapor) is extremely toxic in 
high concentrations. Breathing the vapor may cause severe damage to the nose, 
throat, respiratory passages, and lungs. In contact with the skin or eyes, 
nitric acid causes severe burning; in the case of the eyes, possible destruction 
of the eye tissues can occur with subsequent blindness. The oxidizer also 
vigorously attacks nnost metal and organic nnaterials and in concentrated forin 
may cause spontaneous ignition and vigorous burning when in contact with 
organic materials. 

Physical Properties . IRFNA has a boiling point of 142 F and a freezing 
point of -65°F and is extremely corrosive. The oxidizer is hypergolic when 
combined with UDMH. Specific gravity of this acid is shown in Fig. 3-1. 



3-3 



LOCKHEED MISSILES & SPACE COMPANY 



in*'«'lfP".WWI!:!!'l7"i,'ll.J','V ' l'.,WJ!!IRWI')',' 



jtmt flt»v^w^^ > ^ _ m.yV ' 



IP ^y y ^ -i'^, "i 'W-'T»K^';jP!»»ffT»'t?>--yyg>»;'' j, iJij; 



LMSC-A604141 



3. 2 OTHER MEDIUMS 

In addition to propellants, other mediums used by the Prinnary Propulsion 
System (PPS) are helium for propellant tank pressurization, nitrogen for 
start-tank pressurization, and a nitrogen-freon mixture for oxidizer-pump 
lip-seal pressurization. 

3. 2. 1 Helium 

The helium used for propellant tank pressurization conforn:is to U.S. Bureau 
of Mines Grade A specification. Moisture content of the gas introduced into 
the propulsion system is less than 26. 3 parts per million (ppnn) by volume at 
standard conditions. Cleanliness requirements are shown in Table 3-1. 

3. 2. 2 Nitrogen 

Nitrogen used for pressurizing the start tank bellows conforms to Federal 
Standard BB-N-411, Type I, Class I, Grade B. The moisture content is less 
than 26. 3 ppm by volume under standard conditions. The cleanliness require- 
ments are presented in Table 3-1. 

Table 3-1 
CLEANLINESS REQUIREMENTS OF PNEUMATIC FLUID MEDIA 



Particle Size 
(nnicrons) 

26 to 50 

51 to 100 

101 to 150 

Over 150 



Number Permitted 
Metallic Non-Metallic 



100 

30 







1000 

300 

50 





3-4 



I 



LOCKHEED MISSILES & SPACE COMPANY *■ 



■■•;.t!iK,^^^-^k ■-i'i-!~5tj***ii-«i-./,-v*-i»_'^„.. , 



LMSC-A604l4i 



3.2.3 Nitrogen Freon- 14 Mixture 

The nitrogen freon- 14 mixture used for pressurizing the lip seal is supplied 
from a small regulator through a normally closed-solenoid valve in the 
vehicle attitude -control system. This solenoid valve is energized to open 
any time that liquid is present in the Model 8247 engine system. 

At AMR, the gas is purchased as a mixture with the two gases meeting the 
following specifications: nitrogen conforming to Federal Standard BB-N-411 
Type I, Class I, Grade B; freon conforming to MS-3 1-4000, tetrafluoromethane 
(freon- 14). The mixture has a moisture content less than 20 ppm by volume 
under standard atmospheric conditions. Cleanliness requirements of the 
mixture are shown in Table 3-1. 



3-5 



I Or'l^uc-crr^ KAi 



•II r— *^> 



LMSC-A604141 



3.3 PROPELLANT LOADING 

3. 3. 1 Method of Propellant Loading 

Propellant tanks are loaded with fuel and oxidizer during the final hours before 
vehicle launching. The propellants are pumped through the launch pad 
propellant-transfer equipment and fill lines (in the umbilical nriast) to the 
ground quick-disconnect couplings. These ground couplings attach to the air- 
borne halves of the vehicle fuel and oxidizer fill couplings. Vehicle propellant 
tanks are also equipped with vent lines that terminate at airborne quick- 
disconnect couplings mounted near the forward end of the propellant tank 
assenribly. The vehicle vents are coupled to return lines so that a closed- 
loop circuit is fornned by the propellant-transfer equipment, load lines, 
propellant tanks, and vent lines. By means of this closed-loop arrangement, 
the normal process of venting the tanks during propellant loading can be 
reversed and, by means of pressurization, drain the tanks of propellants, 
if necessary. 

3. 3. 2 Propellant Loading Calculations (Main Tanks) 

The loading of the main propellant tanks for flight is computed as outlined 
in this section. 

Tables 3-2 and 3-3 present the breakdown of non-innpulse and transient- 
impulse propellants for the burns as indicated. Residual, or trapped non- 
innpulse propellants, are shown in Table 3-4. The nonninal propellant tank 
volunne available for propellant loading is shown in Table 3-5. 

Loading calculations are based on actual calibrated tank volunnes for each 

vehicle considered; the vehicle log book contains the calibrated volunnes for 

the particular vehicle. Nominal tank volumes (including sumps) are 77. 45 ft 

3 
for fuel and 100. 50 ft for oxidizer. 



3-6 



I 



LOCKHEED MISSILES & SPACE COMPANY 

"'f"""V"lT"V.WN'.l. I. III . I I i .UJl- I . 1 .I..I, III . i H B " t — v'.- '-" ~. i i ,^ . Ill I, . Ill I . i .-" r""-, /"TT" — ~ '' . ' -.J i J i J '- I JU ii -U ' "i.^r 'i' r^.M«JTMl^-^^ ''' l-' ' >^- '^''.-.' » »W'"w.i' yvJt «fiv-i'»,-iip^^^^ 



LM3C-A604141 



Table 3-2 
TRANSIENT- AND NON-IMPULSE PROPELLANTS INITIAL BURNS 

T^ ,_ Oxidizer 

Non-Impulse Propellants /•,, x 



Pre-flow 

Propellant flow past injector prior to 

thrust-channber ignition 0.00 Z. 00 

Cooling Passages 

Propellant expelled after each shutdown 11. 50 0.00 

'ropellant Isolation Valves 

Propellant expelled after each shutdown 8.69 3.93 

Propellant flow from fuel valve closure 

to oxidizer valve closure 19. 80 0, 00 

Total 39.99 5.93 

Transient-Impulse Propellants 

Propellant flow from to 75% P 1. 58 0. 56 

^ c 

Residual propellant in start tanks 0. 17 0. 86 



Propellant flow from shutdown signal to 

fuel-valve closure 1. 15 0' 45 

Total 2.90 1.87 



1. 


58 


0. 


17 


1. 


15 



Total transient- and non-impulse pro- 
pellants for each initial burn 42. 89 = A 7. 80 = B 



3-7 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Table 3-3 

TRANSIENT- AND NON-IMPULSE PROPELLANTS FINAL BURN 
(Assuming Propellant Depletion) 



Non-Impulse Propellant s 



Oxidizer 
(lb) 



Fuel 
(lb) 



Pre-flow 

Propellant flow past injector prior to thrust 
channber ignition 0. 00 

Cooling Passages 

Propellant expelled after final shutdown 10. 50 

Total 10. 50 



2.00 

0.00 
2.00 



Transient-Impulse Propellants 

Residual propellant in start tanks 
Propellant flow from to 75% P 
Propellant flow from 70 to 0% P 



Total 



0. 17 


0.86 


1.58 


0.56 


8. 20 


0.0 



9.95 



1.42 



Total transient- and non-impulse pro- 
pellants for final burn 



20.45 



3.42 = D 



3-8 



LOCKHEED MISSILES & SPACE COMPANY 



t i , i .i,,i>aaiw^g ) ili|i|i»l<iijljMi i ii]>jf|j.,u. i JW^ 



!»!»eg»i^?«-'%l>«.?;m» ~j.Bliyiy.>.^>:Jiy. i l|^ 



LMSC-A604141 



Table 3-4 
RESIDUALS (TRAPPED NON-IMPULSE PROPELLANTS) 



m 



Feed lines (including PIV's) 

Propellant vapor (tanks at 60 F and 
10 to 12 psia) 

Fill lines 

Fuel stand pipe 

Punnp to engine propellant valves 

Fuel bias 



Oxidizer 
(lb) 

2.3 

1.8 
1.5 
0.0 
5.2 
0.0 



Fuel 
(lb) 

1.6 



1. 


2 


0. 


8 


0. 


9 


1. 


7 


12. 






10.8 = E 



18. 2 = F 



Table 3-5 
NOMINAL TANK VOLUMES 



1^ 







Oxidizer 
(ft3) 


Fuel 
(ft^) 


Nonninal Tank Volumes -:= 




100. 50 


77.45 


Feed, fill, and engine 




0.09 


0.08 


Tank stretch 




0.05 


0.08 




Total. Volume 


100.64 


77.61 


Minimum ullage 




-0.90 


-0.75 



Nominal volume available for propellant 
loading 



99.74 



76.86 



*These are nominal values. Actual tank volumes on individual vehicles will 
be accurately calibrated. 



i 
i 

e 



3-9 



LOCKHEED MISSILES & SPACE COMPANY 



_,.,-— ■..iv.-—*»-"*f^.-W'JMr'^ 



LMSC-A604141 



The fuel tank is loaded first to the minimum ullage value and then, using the 
altitude-corrected engine-systenn-mixture ratio, the oxidizer tank load is 
calculated. The oxidizer tank is checked to verify that loading does not ex- 
ceed the minimum ullage value for the tank. If the loading into the oxidizer 
tank exceeds this nninimum ullage value, the process is reversed; the oxi- 
dizer tank is loaded first, and then, using the corrected nnixture ratio, the 
fuel tank load is calculated. 

Loading of either tank naust be predicated upon the mission profile which 
indicates the number of PPS burns to be attempted. After each shutdown of 
the Model 8247 engine, varying quantities of oxidizer are vented, and due 
to the dissimilarity of these amounts, the loading must necessarily vary 
with the number of burns to be attempted. In the follovs,'ing example arbitrarily 
chosen, propellant tank loading equations and typical calculations for a mission 
profile of 5 PPS burns are presented. 

Fuel Tank Loading Equations 

a. Equation 1 



V = V - V 
fl f uf 



where: 



V 
V, 



fl 



uf 



= volume of fuel to be loaded (ft ) 

3 
= calibrated volume of fuel tank (ft — obtained from 

vehicle log book and Table 3-5) 

3 
= minimum fuel tank ullage. Constant 0. 75 ft 



Note: This equation produces the same values as Table 3-5 



3-10 



LOCKHEED MISSILES & SPACE COMPANY 



b. Equation 2 



where: 



LMSC-A604141 



^fi = ^V^^fJ 



W 



fl 
fn 



weight of fuel to be loaded 

density of fuel at 60°F = 49. 582 Ib/ft^ 



Oxidizer Tank Loading Equations 



a. Equation 3 



where: 



w = W - C 
^fi fl ^1 



W 



fi 



weight of impulse fuel (lb) 

constant total weight of non-impulse fuel, defined as follows; 



where: 



C^ = (M-l) (B) + D + F 



M = number of engine burns in miission 

B = transient- and non-imipulse propellants for 
initial burns (Table 3-2) 

D = transient- and non-impulse propellants for 
final burn (Table 3-3) 

F = residual propellants (Table 3-4) 



3-11 



LOCKHEED MISSILES & SPACE COMPANY 



b. Equation 4 



LMSC-A604141 



W . = (W^.) (R ) 

Ol 11 oo 



where: 



W 



Ol 



R_ 



oo 



= weight of impulse oxidizer (lb) 

= total engine system mixture ratio at altitude. 
Sample calculation shown in Paragraph 6.8. 



c. Equation 5 



where: 



where: 



ol Ol c 



W 



ol 



weight of oxidizer to be loaded (lb) 

constant, total weight of non-impulse oxidizer 



defined as follows: 



(M-1) (A) + C + E 



A = transient- and non-impulse propellants for initial 
burns (Table 3-2) 

C = transient- and non-impulse propellants for final burn 
(Table 3-3) 

E = residual propellants (Table 3-4) 



Check Oxidizer Tank Ullage 



a. Equation 6 



<*ol' 



ol 



on 



3-12 



LOCKHEED MISSILES & SPACE COMPANY 



'^,l ! fMMRy ? !^!l*--. '^'??^'yf'«?:^T^^?r ^*^ Jf^ -y^T'S ^ 'g ^ s■y, ' . l^Hw^yg^'^^■-■y^°^r•-y J ^ ^ '*;^ \'.; '^ ^ . - - 'y^ 



rry-??3i»'gg-*t'r--^-'-T';;>vpiy^^ ■ ' ^ '^ ^ ^^^ f* ??jf pac5r? f^yy >»■' •' '^ ^ V 1 J ^Mftsyi^M- w i ffn rw vi*.'yv'<-*'':;^^s f'g-'^^ 



LMSC-A604141 



where: 

p = density of oxidizer at 60°F = 97. 916 Ib/ft^ 

on ' 

V , = volume of oxidizer loaded 
oi 



b. Equation 7 



V = V . V 

uo o ol 



V = calibrated volume of oxidizer tank (ft — obtained 

from vehicle log book and Table 3-5) 

V = minimum oxidizer tank ullage. Constant - 0. 90 ft' 



3 
V = 0. 90 ft for correct loading 



Example Calculation 



V^ = 77, 21 (vehicle log book and Table 3-5) 

V^ = 100. 64 (vehicle log book and Table 3-5) 

R = 2.551 

M = 5 



Fuel 



Vfi = V^ - V^f (Eq. 1) 

Y^^ = 77. 21 - 0.75 

V^j = 76. 46 ft^ 

^fl = ^fl^^fJ <Eq- 2) 

W^j = (76.46) (49.582) 

Wj^ = 3791.04 1b 



3-13 



LOCKHEED MISSILES & SPACE COMPANY 



Oxidizer 



LM3C-A604141 



W 



fi 



W, 



W 



fi 



= W 



fl ""1 
3791.04 - 

3738. 22 lb 



[(4) (7.80) 



+ 3.42 + 18. 20 



] 



(Eq. 3) 



W . 
oi 

W . 
oi 

W . 
oi 



(3738.22) (2.551) 
9536. 20 lb 



(Eq. 4) 



W 



ol 



Ol 2 



W 



ol 



= 9536.20 + [(4) (42.89) + 20.45 + 10. 8o] 



W 



ol 



= 9739.01 lb 



(Eq. 5) 



Ullage Check — Oxidizer 

W 



V 



ol 



ol 



V 



ol 



ol 



on 



9739.01 
" 97.916 

= 99.46 ff^ 



(Eq. 6) 



V = V 

uo o 



V 



uo 



V 



uo 



V 



ol 



100.64 - 99.46 
1. 18 ft^ 



(Eq. 7) 



The actual ullage volume is greater than 0.90 ft ; therefore, the loading is 
acceptable. 



3-14 



LOCKHEED MISSILES & SPACE COMPANY 



' ^ 'i y ' i)^ 'i l i|i jiyj l ^pi_,^pi,)4iu|» ii ii i jijtt!jiBy_tiiy^ i i^ , «fim! s j Kmi^M .i fu i ±i Mr^'r ^»4^ f m._.ii r j, ' .xA ' ? r7^--''^ 



LMSC-A604141 



In the aforementioned example, in which fuel was loaded first, V^^ was 

acceptable. Frequently, due to variations in tank volunnes and inixture ratio, 

V , will not be sufficient when calculated by this method. In such cases, it 

uo 
is necessary to load the oxidizer first, substituting oxidizer parameters for 

fuel parameters in the existing equations. The following is an example of 

such a condition: 



Example Calculation: 



Fuel: 



^f 


= 


77. 61 


V 
o 


= 


100.64 


R 


= 


2. 551 


M 


— 


5 



(vehicle log book and Table 3-5) 



V 



fl 



V 



fl 



V 



fl 



= V 



V 



f 'uf 
77.61 - 0.75 

76.86 ft^ 



(Eq. 1) 



W 



fl 



W 



fl 



W 



fl 



(76.86) (49. 582) 
3810.87 lb 



(Eq. 2) 



3-15 



LOCKHEED MISSILES & SPACE COMPANY 



Oxidizer: 



LM3C-A604141 



*fl 


~ 


^fl - ^1 


''{i 


= 


3810,87 - [(4) (7 


*£i 


= 


3758.05 lb 


w . 

Ol 


= 


W^. (R ) 

fl OD 


W . 

Ol 


= 


(3758.05) (2.551) 


W . 

Ol 


= 


9586.78 lb 


ol 


= 


oi 2 



w 



ol 



W 



(Eq. 3) 



(Eq. 4) 



(Eq. 5) 



ol 



9586.78 + [(4) (42.89) + 20.45 + 10. 8o] 
9789. 59 lb 



Ullage Check — Oxidizer: 

W 



ol 



ol 



(Eq. 6) 



on 



V 



ol 



V 



ol 



9789. 59 
" 97.916 

= 99.98 ft' 



V = V 

uo o 



ol 



(Eq. 7) 



V = 100.64 - 99.98 
uo 

V = 0.66ft^ 
uo 



The actual ullage volume is less than 0.90 ft ; therefore, the loading is un- 
acceptable. 



3-16 



LOCKHEED MISSILES & SPACE COMPANY 



''i'BWftl-*K4I'-"-'f!!!(<H»TO»ii!!^IH?!WW 



^.i3';!»''^^''j» ;y '^ ;B ' ?Vi '' - '" .ww'8*y',yvpi P !(twt f iH#^ 



s^tawTweKp*!?^^. 



LMSC-A604141 



It is necessary to recalculate the loading, starting with the oxidizer. The 
procedure is as follows: 



Oxidizer: 



Fuel: 



V 



ol 



V 



ol 



V 



ol 



= V - V 

o uo 

= 100.64 - 0.90 
= 99.74ft-^ 



(Eq. 1) 



W 



W 



W 



ol 



ol 



ol 



(99.74) (97.916) 
9766. 14 lb 



(Eq. 2) 



I 



I 
I 
I 



W . 

Ol 

W . 

Ol 

W . 

Ol 



^if 



W. 



W 



if 



W 



fl 



W 



fl 



W 



fl 



= W 



ol 



(Eq. 3) 



9766. 14 -[(4) (42.89) + 20.45 + 10. 8o] 
9563.33 lb 



W 



R 
oo 

9563. 33 

2. 551 

3748.86 lb 



W., + C^ 



3748.86 + [(4) (7. 80) + 3. 42 + 18. 2o] 
3801.68 lb 



3-17 



(Eq. 4) 



(Eq. 5) 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Ullage Check - Fuel: 



W 



V 



fl 



fl 



V 



fl 



fl 



V 



"fn 




3801.68 


■ 49. 


582 


: 76. 


67 ft^ 


= ^f 


- ^fl 



(Eq. 6) 



uf 



V ^ = 77.61 - 76.67 
u f 

V , = 0.94 ft^ 
uf 



(Eq. 7) 



The actual fuel ullage volume is greater than 0. 75 ft ; therefore, the loading 
is acceptable. 

3. 3. 3 Start Tank Charging 

Servicing of the Model 8247 engine start system is accomplished with two 
AGE servicing carts. These carts supply propellants filtered to 25 microns 
absolute and start tank charging N^-gas filtered to 10 microns absolute. 

Due to the critical nature of start tank loading, the method of loading is 
presented below: 

3.3.3.1 Fuel Start Tank Loading 

Servicing the Fuel Loading Cart . The fuel start tank loading cart is filled 
from the launch pad fuel bulk stores. A minimum of 130 in. of UDMH is 
loaded into the cart. A sample of the fuel is then taken from the system and 
analyzed for conformity to MIL-D-25604B- 1. The oxidizer and fuel nitrogen- 
charging bottles located in the fuel loading cart are pressurized to approxi- 
mately 1500 psig. 



3-18 



LOCKHEED MISSILES & SPACE COMPANY ■" 



m 



LMSC-A604141 



Fuel Start Tank Loading. The gas side of the start tank is pressurized 
(Item 1, Fig. 3-2) to stack the bellows, and the start tank-mounted liquid 
bleed valve (Item 1, Fig. 3-3) is opened - thereby forcing out trapped gas 
or liquid from the liquid side of the tank. The bleed valve is closed and the 
AGE Hquid-transfer line is connected to the start tank fill valve (Item 2, 
Fig. 3-3). The volume to be transferred to the start tank is variable and 
must be determined from Table 3-6. 



Table 3-6 
QUANTITY OF UDMH TO BE LOADED INTO FUEL START TANK 



Fuel 
Tennperature 
(^F) 

40 ± 5 

50 ± 5 

60 ± 5 

70 ± 5 

80 ± 5 

90 ± 5 
100 ± 5 
110 ± 5 



Dry Load* 

Volume 

(in. 3) 

106. 1 ± 0.9 
106.9 ± 0.9 
107.7 ± 0.9 
108. 5 ± 0.9 
109.3 ± 0.9 
110.0 ± 0.9 
111.0 ± 0. 9 
111.8 ± 0.9 



Wet Load** 


Volume 




[in. 3) 


93 


c + 2. 
■5 - 


94. 


^ - 


95. 


„.z.o 


95. 


3.Z.0 


96. 


A + 2.0 
^ - 


97. 


4 + 2.0 
- 


98. 


'^ - 


99. 


, + 2. 



■A fuel loading for a dry system includes the fuel necessary to completely 
fil an empty start tank bellows, the fuel system lines (to Ihe gas generSor 

_ solenoid valve and dual check valve), and instrumentation. generator 

"■•A fuel loading for a wet system does not include the start tank residual, 
fuel system lines, and instrumentation volumes. 



3-19 



LOCKHEED MISSILES & SPACE COMPANY 



'^^.•Wi^Wr TH'i'^^frdiPJ^ . ,■ 



LMSC-A604141 




1 Fuel Start Tanlc Nitrogen Fill Port FP-11 

2 Oxidizer Start Tank Nitrogen Fill Port FP-10 



Fig, 3-2 Fuel and Oxidizer Start Tank Nitrogen Fill Ports 



I 
I 



3-20 



LOCKHEED MISSILES & SPACE COMPANY 



J 



TO9-»^TT-.-,-iKr«r«i twawsr,^?rwsw«?r'" 



*,^^, p^pyr^^- ,A-.,; L>i!), i y^ l^ h; Pl^^^ ^l WJSg■rr:yp^^|^W|g j^ i |V g.■^ ll . jW b^lJ ,. ^ J g? 'V:^y?'^'- ' ^ '**^^'^^^^^y^^?^^1^^^ 



LMSC-A604141 




1 Fuel Start Tank Valve Bleed Point BP-4 

2 Fuel Start Tank Fill Port FP-6 

3 Fuel Dual Check Valve Outlet TP-14 

4 Fuel Dual Check Valve Inlet TP-16 

5 Oxidizer Start Tank Valve Bleed Point BP-5 

6 Oxidizer Start Tank Fill Port FP-9 



Fig. 3-3 Engine Start Tank System 



3-21 



LOCKHEED MISSILES & SPACE COMPANY 



■■iyi 



LMSC-A604141 



3. 3. 3. 2 Oxidizer Start Tank Loading 

Servicing the Oxidizer Loading Cart . The oxidizer start tank loading cart 

3 
is filled from the launch pad oxidizer bulk stores. A minimum of 35 in. of 

oxidizer is drawn into the sight glass, and all AGE plumbing is bled in order 

to elinninate the possibility of loading trapped voids into the engine start tank. 

A sannple of the oxidizer is then taken from the system and analyzed for 

conformity to MIL-P-7254-1 Type IIIB. 

Oxidizer Start Tank Loading . The gas side of the start tank is pressurized 
(Item 2, Fig. 3-2) to stack the bellows, and the start tank-nnounted liquid- 
bleed valve (Item 5, Rg. 3-3) is opened — thereby forcing out trapped gas or 
liquid from the liquid side of the tank. The bleed valve is closed and the AGE 
liquid-transfer line is connected to the start tank fill valve (Item 6, Fig. 3-3). 
The volume to be transferred to the start tank is variable and nnust be deter- 
mined from Table 3-7. 



3. 3. 3. 3 Nitrogen Precharging of Start Tanks 

Fuel Tank . The fuel nitrogen-charge tank (located on the fuel start tank 
service cart) is charged with gas at 1480 to 1500 psig. This charging tank 
is connected to the start tank at FP-11 (Item 1, Fig. 3-2). Using the UDMH 
temperature in the start tank and the gas temperature in the nitrogen charge 
tank, the proper charging pressure is determined from the appropriate curve 
(Figs. 3-4a through 3-4h). The charging tank is bled slowly (so as not to 
reduce gas temperature and change charging requirements) to pressure "A", 
pressurizing the start tank bellows to the limits of pressure "B". The gas 
charging line is now removed, and all ports are pressure capped. 

Oxidizer Tank . The oxidizer nitrogen-charge tank (located on the fuel start 
tank service cart) is charged with gas at 1325 to 1350 psig. This charging 



3-22 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



tank is connected to the start tank at FP- 10 (Item 2, Fig. 3-2). Using the 
liquid temperature in the start tank and the gas temperature in the nitrogen 
charge tank, the proper charging pressure is determined from the appropriate 
curve (Figs. 3-5a through 3-5h). The charging tank is bled slowly (so as not 
to reduce gas temperature and change charging requirements) to pressure 
"A", charging the start tanks to the limits of pressure "B". The gas 
charging line is now removed, and all ports are pressure capped. 



Table 3-7 
QUANTITY OF IRFNA TO BE LOADED INTO OXIDIZER-START TANK 



Oxidizer 

Temperature 

(°F) 

40 ± 5 

50 ± 5 

60 ± 5 

70 ± 5 

80 ± 5 

90 ± 5 
100 ± 5 
110 ± 5 



Dry Load* 

Volume 

(in. 3) 

17. 5 ± 0. 16 

17.6 ± 0. 16 

17.7 ± 0. 16 

17.8 ± 0. 16 

17.9 ± 0. 16 
18.0 ± 0. 16 

18. 1 ± 0. 16 
18. 2 ± 0. 16 



Wet Load** 


Volume 
(in. 3) 


13.6 


+ 0.5 
- 


13.7 


+ 0.5 
- 


13.8 


+ 0.5 
- 


13.9 


+ 0. 5 
- 


14.0 


+ 0.5 
- 


14. 1 


+ 0.5 
- 


14.2 


+ 0.5 
- 


14.3 


+ 0.5 

A 



*An oxidizer loading for a dry system includes the oxidizer necessary to 
completely fill an empty start tank bellows, the oxidizer system lines (to 
the gas generator solenoid valve and dual check valve), and instrumentation. 
**An oxidizer loading for a wet system does not include the start tank residual, 
oxidizer system lines, and instrumentation volumes. 



3-23 



LOCKHEED MISSILES & SPACE COMPANY 



yt J W H.W ' W. 'i . 'i" i!:ilJ 



LMSC-A604141 



900 













^ 




/ 


■ CART 
PRESSURE^ 




-^ 


« fTART 










^ 


^^^ TANK 
■^^ 4 LIMITS 


^ 


."■^"^^ ^ 




^^ 






1^ 













40 60 80 

TEMPERATURE (°F) 



Fig. 3 -4a Fuel To Be Loaded at 40° iS'F 













^ 




A -CART 
PRESSU 


""V- 


^ 


^ 


IT- — B . 
\A START 






^ 


^^"^ -- 




^^^-^ TANK 
\, LIMITS 



























40 <0 80 

TEMPERATURE <°F) 



Fig. 3-4b Fuel To Be Loaded at 50° ±5°F 



3-24 
LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




40 60 80 

TEMPERATURE (°F) 



Fig. 3-4c Fuel To Be Loaded at 60° ±5°F 



B 
I 
I 




40 '0 80 

TEMPERATURE (°F) 



Fig. 3-4d Fuel To Be Loaded at 70° ±5°F 



3-25 



LOCKHEED MISSILES & SPACE COMPANY 



frTryp>^^^rr>tri^' 



LMSC-A604141 





A . CART 

PRESSURE — ^ 




. 


/ 






^ 


y^ 


y^ 


K START 
VS. TANK 
^^ LIMITS 






^ 


^ 


^ 




^ 




^^ 








'^^^' 













40 60 ' 80 

TEMPERATURE (°F) 



Fig. 3-4e Fuel To Be Loaded at 80' ±5'F 



i3aD 
1200 




A - CART 








^ 












1100 
1000 
900 
800 






i> 






\\ START 

< TANK 

y^ i LIMITS 














^ — ^ 




^^^ 





















40 60 80 

TEMPERATURE (°F) 



Fig. 3-4f Fuel To Be Loaded at 90° ±5°F 



3-26 



LOCKHEED MISSILES & SPACE COMPANY 



!?!gjjjlgjyi ^ ' »J!) i jij^^ l . l| )W ^ |jyjf..a^ 



LMSC-A604 141 



1300 














/ 






* -CART 














PRESSURE — > 


\ 


^ 






IIMJU 








\ 




^■"^ ^ 


rV START 
^^ TANK 
i, LIMITS 






^ 


^y^ ^ 








1000 




^ 










VUU 
nnn 


^ 













40 60 80 

TEMPERATURE (°F) 



100 120 



Fig. 3-4g Fuel To Be Loaded at 100° ±5°F 





A . C 
P 


IkRT 
RESSURE — K 


^ 




y^ 




^ 


\ 


^ 






^^ START 
.' ^V TANK 
^ LIMITS 


^ 


^ 


^ 


1^ 


^ 




/^ 

























40 60 eo 

TEMPERATURE (°F) 



100 120 



Fig. 3-4h Fuel To Be Loaded at 110° ±5°F 



3-27 



LOCKHEED MISSILES & SPACE COMPANY 






g^jaa?r ,fcjwiMJ^t ' rB g a i f^ ' ff s g r:3gy?yf^ 



LMSC-A604141 



















A-C/kRl 




^ 


/ 






PRESSURE — V 


^ 


START >\ 
TANK ^ 
■ LIMITS ^ 




^ 




^ 






^ 























TEMPERATURE ( P) 



Fig. 3-5a Oxidizer To Be Loaded at 40° ±5°F 




40 60 eo 

TEMPERATURE (°F) 



Fig. 3-5b Oxidizer To Be Loaded at 50° ±5°F 



3-28 



*? iW;,l»W!^*;*W''Wf> ' - '' W.j.*. ^ Bt B {8 ! %' B i'*^ ~.' ' ^.^ ^^^^ ^ ^ ^ 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




40 60 80 

TEMPERATURE (°F) 



1300 














liOO 














^ 


^ 

% 
















\ 


3- START 


1100 




A 


■CART 






^,-. — ' 


\ TANK 




^ 


^ 




>^ 


i::^ 


^ 




1000 
900 


A 


y' 


^ 









800 


2C 


4 


) 


6 


0-— 8 





lo 


Is — 


1i 



TEMPERATURE ( F) 



Fig. 3-5d Oxidizer To Be Loaded at 70° ±5°F 



3-29 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




40 60 

TEMPERATURE ("fl 



Fig. 3-5e Oxidizer To Be Loaded at 80° ±5°F 

















A . CAR 
PR! 


T 

SSURE — 1 




^ 


W START 
Jr TANK 
"^ \ LIMITS 






L> 


:^ 


^ 




^ 


^ 


^^ 








^''^^^^ 













40 (0 SO 

TEMPERATURE (°F) 



Fig. 3-5f Oxidizer To Be Loaded at 90° ±5°F 



3-30 



LOCKHEED MISSILES & SPACE COMPANY 



■^w^wwiMHiB '»iij,i L! j> I \ . ...<aisKmim>v<f«.'rw 



'■ '-?^'T»^K'i^e«''!S!«r^WaK5!51MWWr^^SE?T 



jirr^TjiniBWSWf.^ 



LMSC-A604141 



1300 




40 iO 80 

TEMPERATURE (°F) 



100 120 



Fig. 3-5g Oxidizer To Be Loaded at 100° ±5°F 













^ 




A -CA 

PR 


RT 

E55URE 1 


. 




y^ 

V)l START 
^>f^ TANK 
"^ \ LIMITS 








y^ 


^ ^-^ 




^^^"'^ 




y^' ^^ 





















40 60 

TEMPERATURE (°F) 



Fig, 3-5h Oxidizer To Be Loaded at 110° ±5°F 



3-31 



LOCKHEED MISSILES & SPACE COMPANY 



r*?l ^^^jMff f ayi')yy i»s-y»?y^5^?>wr^ : '' ;? ,?g J ^^!^ fr ' ^'^y!W^ty ? l > ILr; "-; ' '^. ' L,^jW' W .1 !| f fg.'-. "^■■W"!?!^ '?^ 



LMSC-A604141 



Section 4 
ENGINE OPERATION 

When propellants are supplied to the turbopunnp inlets at suitable pressures, 
engine operation is initiated by applying a Z8-vdc starting signal. However, 
during the coast period prior to engine start, the lack of gravitational force 
would pernnit the gaseous ullage to enter the propellant feed lines. If a 
system was not provided to prevent gas fronn entering the feed lines during 
engine starting (includes initial start and subsequent restarts), then inter- 
rupted burning and possible engine malfunction could occur. 

4. 1 ULLAGE ORIENTATION CONTROL 

During coast, a vehicle in a relatively low orbit, such as the Gemini-Agena 
Target Vehicle (ATV), is in a "near zero, relative g" field. The centrifugal 
force of the vehicle rotating around the earth and the earth's gravitational 
force cancel out each other, leaving as the only unbalanced force the slight 
deceleration (approximately iC^g) due to aerodynamic drag forces on the 
vehicle. Since the Gemini ATV is a "forward flying" satellite, the propel- 
lants seek equilibrium positions in the forward ends of the respective tanks. 

After a booster-engine cutoff or an Agena main-engine shutdown, the 
residualpropellants seek to become oriented at the forward end of the pro- 
pellant tanks. In a "zero g" field, wetting propellants tend to climb the tank 
walls, thus increasing the liquid-gas interface and reducing the unwetted 
area. For small ullages or large amounts of fluid, the liquid may cover the 
walls completely and displace the ullage to the center of the propellant tank, 
thus indicating ullage orientation prior to Agena first-burn may not be re- 
quired. In order to ensure Primiary Propulsion System (PPS) start or 



4-1 



LOCKHEED MISSILES & SPACE COMPANY i 



LMSC-A604141 



restart, propellant must fill the engine and feed lines and the containment 
"sumps" and cover tank outlets to a height sufficient to ensure that the 
pumps do not suck a hole through the liquid and ingest a prohibitive amount 
of pressurization gas. 

The containment system, described in Paragraph 2. 3. 2, is designed to trap 
a sufficient quantity of propellants over tank outlets, thus allowing start of 
and maintaining combustion in the main engine until engine thrust has 
oriented the bulk propellants over the tank outlets. However, because of 
multiple burn requirements of the ATV mission, Agena ullage orientation 
is to be accomplished by an artificial gravity field produced by the 16 lb- 
thrust units of the Secondary Propulsion System (SPS). 

A study has been made to determine the thrust-time relationship required 
to properly orient propellants for engine start at thrust levels between 16 
and 400 lb, and for vehicle masses between 2, 000 and 16, 000 lb. Results 
of this study show that for a given vehicle weight, the propellant orientation 
period is approximately inversely proportional to the square root of the 
applied thrust. Figure 4-1 is used for determining the length of the propel- 
lant orientation period (defined as the time from SPS ullage ignition to the 
time when a propellant head above pump inlets is obtained which is sufficient 
to ensure reliable main engine start). For the preliminary mission profile, 
with assumed propellant-orientation burn times of 74 and 23 sec for the 16 
and 200 lb-thrust units, respectively, the 16 lb-thrust unit leaves approxi- 
mately 63, 400 lb/sec of impulse propellant for orbital rendezvous maneuvers; 
while the 200 lb-thrust unit leaves approximately 15, 600 lb/sec of impulse 
propellant. (Note: values are approximate and are presented to illustrate 
relative quantities. ) 

Two ullage orientation periods are provided in which the 16 lb-thrust units are 
used. Each period is of sufficient duration to provide propellant orientation 
with only one unit functioning at the lower thrust limit (14. 4 lb). The short 



4-2 



LOCKHEED MISSILES & SPACE COMPANY * | 



LMSC-A604141 



o 
o 



























-1 








1 


Ij 




A 


LLI 








/ 








- £ ^ - 

■o Z 

— O LU — 








/ 








^ 2 aJ 
:5 VI K* 








/ 






J3 
O 

o 

CN 


16 lbf<F<40( 

2,0 00 lbm<Wv 

NCLUDES A 

PERIOD 








/ J 


1 




' 










/ 


















/ 


/ 






i I 








/ 


/ 


/ 














/ 




/ 






3 

^0 


Q 
u 

c 

> 

r 


J 

< 

1 


/ 


/ 


_j / 

UJ / 

r) / 

U- / 


/ 






' 


I 


y 




/ 


/ 

























Di 

o 





2 




O 


<N 


t- 


O 


< 


X 


d: 


'— 


LU 




_i 


00 


UJ 




u 




u 




< 












lO 




UJ 






O 



o 




bX) 
I— ( 

OJO 



o 


d 




n 


M 


• r-» 


n 


+J 


•i-i 


n) 


u 


f-i 


(h 


.— 1 


Ti 


0) 


O 


u 


• ■-< 


u 




< 


^ 


r-l 


d 




u 


42 


•4-> 


u; 




> 


d 


tn 





tn 


M 


<U 


o 


d 




o 




d 


1— t 


(U 


0) 


e 




Q 


Ph 


in 


1— < 


?l 


n1 


UJ 


-M 


tH 


O 


<u 


H > 


1— { 




•^ 





60 



(D3S) aOia3d NOIlVlN3iaO lNVT13dOyd 



4-3 



B 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



period (18 sec) is used prior to first burn when thrust/weight is the lowest. 
As this ratio increases, the ullage orientation period increases (Fig. 4-1). 
A 70-sec ullage orientation period is provided for use on all subsequent 
burns. 

4. 1. 1 Secondary Propulsion Systenn Operation 

A brief discussion of the secondary propulsion system is included here, 
because of the use of the SPS in providing PPS ullage orientation control. 
Detailed SPS analysis is presented in the Secondary Propulsion System 
Engineering Analysis Report , LMSC-A387649- 1 • 

The BAG Model 8250 secondary propulsion system (Fig. 4-2) is a storable 
liquid-propellant system which is designed to supply thrust, on demand, for 
multiple firings under vacuum and zero-gravity conditions for a period of 
30 days. This system is completely contained, including positive expulsion 
propellant tanks, gas pressurization system. 16 lb-thrust (Unit 1) thrust- 
chamber assembly, and 200 lb-thrust (Unit II) thrust-chamber assembly. A 
module consists of the components described above, and each vehicle has 
two modules - one located on the +Y axis and the other on the -Y axis. Each 
module is capable of producing a minimum total impulse of 40, 000 lb-sec. 

Components - The SPS fluid flow system (Fig. 4-3) includes the following 
components*: 

(1) Spherical gas-storage tank providing high-pressure N^-gas 
initially at a pressure of 4000 psig to the gas pressurization 
system for expelling the propellants and actuating the propel- 
lant valves. 

(2) "On-off" type solenoid-actuated start valve isolating the high- 
pressure gas from the remainder of the pressurization system. 
An electrical impulse actuates the valve to pressurize the gas 
pressurization system. 

* Paragraph numbers correspond to numbers on Fig. 4-3. 

4-4 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



■SPSUNIT II (200 POUND) 
THRUST CHAM.BLR ASSEMBLY 



t^ 




I ■ 




.8247 

NOZZLE 

EXTENSION- 



SPS 
MODULE 



GEMINI PECULIAR 

AL: ■•LIAKY STRUCTUR 



..:^ 



\ 



\ i 
\1 



/; 



)■ 



r "V; . 






y 



..y 



-^ 






^^4 



''Sb«^^^4S^..^«'^"£^:£^»i^k>au<''^ 



^^^Hi^i^.^i^u^.^tt 



ST D. AG EN A 
AFT t.UL KHEA,D 



- .. .^:?^-..^.;>;^^. .i-...''^^. 




Fig. 4-2 Secondary Propulsion Jvlodule —Aft View 



4-5 



LOCKHEED MISSILES & SPACE COMPANY 



^^'fS^WRWCISSWWMRK' 



fflr-^^^wwwwwwweEsrKsrjrw?!^??!^ •^.?2TT^p;«r«^5»''='^*Tn^>'«'^^^^ >T'^^g???j«w«*«^J!s^^ 



LMSC-A604141 



iJ 



S S 3 3 




♦ 


si 4 


J=ULB-~ 

1 " 


£0 




i 
g J. i 

sg3?s ^KSri ess 


5 

si 






% 

1. 


J 


s 


>- 






m 

>- 


g 
>- 











B 

4:! 
o 
w 

^ 

o 

I— 1 

•H 

g 

(1) 

•tJ 
w 

>> 
o 

(U 

CO 

1—1 

:i 
a, 
o 
u 

u 

nJ 

C 
O 
o 

<u 

CO 

ro 
I 



4-6 



LOCKHEED MISSILES & SPACE COMPANY 



!l ll g"}>SWi^ j ti ^AW ; f '. ' .t ' ««B! >iHi.«i>yB^)WSJ. ij | syj? j fB i; ' ' vi j; i j!iMji g j i iijyii i ji^ j^ wt p '^ i ^ w ^.^^ i ju^ i puqw ';' ! i --*»— » ?:<- jw j jV -? ' ! ^ .i.^hv** ' ^- ! -""' ' " ^* ^"'■'i^- n ' * fy ^? ^^^l^< ' yyf V?: ' lj g -5 -^r^^ * ' ' ' ■.>/ '..^l ^ -^Al- ^ f " v." f Li^ gg^f7^.ig 



LMSC-A604141 



(3) 



(4) 



(5) 



S ngle-stage gas pressure regulator for reducing the source gas 
pressure to an operating pressure of Z05 psia for the ent.re 

pressur..ed.tanl.propellant-feed system, and for propellant 
valve actuation. "penant 

CylindHcal fuel and o.id.zer tanks each containing a bellows 
type n^etallic, n^nlticycle. e.pnls.on deWce which provid J 
pos. we e.pnlsion „nde. .e.o-gravit, conditions. Lopel ants 
a» d.scha.ged f.on. ..e tan.s into the p.opellant feed s 

the Unu I and Unit II th.ust-chan.ber assembUes. 



(6) 



(7) 



(8) 



-ding fo. th.ust-chantbe, inte.changeability. 

'^::ZT ^°'^"°'^-— "«^ .as-actuated p.opellant valves 
provtdtng for sequencing of the fuel and oxidizer to the thrust- 
chamber assemblies. i-nrust 

Three-way solenoid valves venting the propellant valve actuation 
P^essure at shutdown and providing "on-off" control of theTa 
flow uttlt.ed for opentng and closing the propellant valves 
Untt I and Unit H thrust-cha„ber asset^blies for producing thrust 

1. and 1. 15 and chamber pressures of 79. and 
96. ps.a. respectively. Starting is effected by the hypergolic 
Ignition of fuel and oxidizer. ypergoUc 

0£SIiiii2E- The secondary propulsion system is prepared fo 

electrically energising the start valve ,1 sec before o' °''""°" '' 

lant valves. . U-see start valve lead tim a ZZ" ""^ ''"''''- 

:::r r ^vr "-^"-^ -- ^^'°- ^--^-^:^::::^: "^ 

starts. Accordingly, gas is permitted to flow from the .a. 

-ough the gas filter, start valve, gas pressure rriL:: "ZZ^^ 



4-7 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



valve into the gas side of the propellant tanks. The actuation-gas supply- 
lines of the propellant valves are also pressurized. As the pressure down- 
stream of the regulator increases to Z27 psia, the regulator approaches 
lock-up. 

After the pressurization period, the thrust-chamber assem.bly is operated 
by electrically energizing the solenoid propellant valve. Opening the valve 
establishes continuous liquid-propellant flow from the two propellant tanks 
through the propellant valve into the chosen thrust chamber. The gas pres- 
sure regulator maintains an automatically controlled gas pressure on the 
bellows in the fuel and oxidizer to provide positive expulsion of the propel- 
lant at any altitude and zero-gravity conditions. Liquid propellants enter 
the thrust chamber with a very slight MON lead and ignite hypergolically. 
As the flow continues in a steady state with the gas pressure regulator 
maintaining rated pressure, the thrust chamber develops rated thrust. The 
operation cycle is terminated by simultaneously de- energizing the three-way 
solenoid valve and the start valve. 

The SPS complements the Primary Propulsion System by providing thrust 
required to make orbit adjustments, when the required impulse is less than 
that which can be supplied by the PPS, in addition to providing ullage orien- 
tation control. Unit I and Unit II thrust chambers are always operated in 
pairs for balanced thrust. The Unit II thrust chambers are used for orbit 
adjust, and the Unit I thrust chambers are used for ullage orientation. 



4-8 



LOCKHEED MISSILES & SPACE COMPANY " 



LMSC-A604141 



4. 2 ENGINE START SEQUENCE (FLIGHT) 

When propellants are available at the tank outlets at sufficient head to pre- 
vent gas ingestion, the engine is ready for operation. No electrically 
armed condition is required for the Model 8247 engine. 

For the following discussion, refer to the propellant flow schematic 
(Fig. 2-2), and to the wiring diagram (Fig. 2-5). 

At the predeterminfed time during the ullage orientation period, the engine 
start sequence begins with the application of a 28-vdc signal to pins A and 
B of connector J6000, along with a simultaneous grounding of pins D and E. 
The start signal passes through the electronic gate circuitry, which acts 
as a junction box for dispersing the signal to plug P6011. Pins H and Z of 
P6011 furnish power to the gas generator fuel solenoid valve (GGFSV); pins 
J and a furnish power to the gas generator oxidizer solenoid valve (GGOSV); 
and pins K and b supply power to the pilot-operated solenoid valve (POSV). 
The three solenoid valves actuate simultaneously upon receiving the start 
signal. 

Prior to launch, the start tank bellows and lines are loaded with the proper 
volume of propellants. Opening of gas generator (GG) solenoid valves 
initiates start tank operation. The start tank ullages (outside of bellows) 
are charged to the prescribed pressure (approximately 1000 psia). With 
the opening of the solenoid valves, propellants flow into the gas generator 
due to the force of the gas pressure on the bellows. The propellants ignite 
hypergolically in the gas generator which provides pressure buildup and 
sufficient energy to rapidly overcome the turbine-wheel break-out torque, 
resulting in turbine acceleration. As propellant flow to the gas 
generator continues (due to bellows expulsion), the ullage volume in 
the start tanks increases, the pressure decays, and the turbine accelerates 



4-9 



LOCKHEED MISSILES & SPACE COMPANY 



»ir">snvj,!.tni*»i»j'.4)s."W'»'?rrt5?f 



LMSC-A604141 



rapidly simultaneously with propellant combustion in the gas generator. 
Since the pumps are coupled to the turbine wheel, the pump-outlet pressures 
rise consistent with the relatively high horsepower availability. The process 
of start tank pressure decay and pump outlet-pressure increase continues 
until the two pressures equalize, at which time propellant expulsion from 
the start tanks stops, because the pump-outlet pressure is sufficient to 
force propellant back into the tank, thereby compresang the bellows and 
starting the recharging process. 

At this time, the gas generator is being fed with the propellants from 
the pumps and the bootstrap operation has been initiated. With the 
build-up of turbine speed and increase of pump-outlet pressure, the re- 
charging process is maintained until the fuel and oxidizer valves open and 
thrust-chamber ignition occurs. After ignition, steady-state operation is 
achieved, and the start tank propellant supply is replenished and maintained 
by steady pump-discharge pressure. The check valves prevent backflow of 
propellants during orbital coast, thus continuing the flow of propellants 
within the start tanks and assuring subsequent starts. 

At the same time that the GG solenoid valves open, the pilot-operated 
solenoid valve opens for initiating the sequence which shuts off flow through 
the fuel valve, thus allowing actuation pressure to build up as described in 
Paragraph 2, 2. 8. When the actuation pressure reaches 365 psig, the fuel 
valve main-poppet opens. On initial or phase "A" start, fuel flow to the thrust 
chamber does not take place until the main-discharge pressure reaches 
525 psig and ruptures the frangible disc. During subsequent starts, fuel 
flows into the thrust chamber immediately upon opening of the fuel valve. 

Simultaneously with the rise in fuel pump-discharge pressure, the oxidizer 
pump-discharge pressure increases; at 225 to 300 psig, the oxidizer valve 
is opened when a spring-loaded poppet is unseated. As soon as this poppet 



4-10 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



I 



opens, the frangible disc ruptures, since the rupture-pressure >s 180 ps.g. 
oxidizer then Hows through the thrust-chamber cooling passages to the 
oxidizer injector tnanitold. For all practical purposes, initial and sub- 
sequent start operafons are ident.cal, because the frangible disc ruptures 
allst .mn,ed.ately upon opening of the poppet. Inasmuch as the frangible 
disc is only used to prevent leakage of propellant during the booster phase of 
flight, no problem exists. 

Under normal operating cond.tions, oxidizer enters the thrust chamber 
first Then fuel enters through the injector wift two jets of fuel imprngmg 
on each oxidizer jet. These propellants igmte hypergolically to mrt.ate 
thrust-chamber operation. The time from appHcation of start signal to the 
time when 70 percent P, (thrust chamber pressure) rs reached is considered 
the start transient, and rs nominally 1 sec. Steady- state thrust rs defmed 
as the time from 70 percent P., start to P, decay on engine shutdown, 
operating temperatures are below normal, the fuel may enter the thrust 
chamber before the oxidizer without appreciable change in the engrne start 
transients. 

Once steady- state thrust is achieved, the engine continues to operate nom- 
inally as long as the start signal is applied. 



4-11 



LOCKHEED MISSILES & SPACE COMPANY 



"T^vw-irrwr >?f*(^*E^ 



LMSC-A604141 



4. 3 ENGINE SHUTDOWN (NORMAL FLIGHT) 

Removal of the start signal from pins A and B and the ground from pins D 
and E initiates the engine shutdown sequence. Loss of electrical power 
allows the solenoid valves (GGFSV, GGOSV, and POSV) to close under 
spring and dynamic load. 

Closure of the POSV permits fuel valve -actuation pressure to decay, as 
fuel passes through, rather than stagnating and holding the fuel valve open, 
with the resultant shotting off of fuel flow to the thrust chamber and termina- 
tion of combustion within. The closure of the GGFSV and GGOSV shuts off 
flow of fuel and oxidizer, respectively, to the gas generator and terminates 
combustion in the GG. The lack of combustion products causes the turbine 
to decelerate, and accordingly, the fuel and oxidizer pump pressures 
decrease. Decrease of the oxidizer pressure allows the spring-loaded 
oxidizer valve to close. 

4. 3, 1 Engine Shutdown (Turbine Over speed) 

If at any time during engine operation a turbine over speed condition 
(29, 500 ±500 rpm) occurs, the engine shuts down automatically by the 
electronic gate which senses the overspeed condition as described in 
Paragraph 2. 2. 1. The electronic gate then trips the relay contacts K-1 
and K-2 (Fig. 2-6) to the open condition, thus interrupting the 28-vdc 
signal to the solenoid valves. Removal of electrical power allows the 
solenoid valves (GGFSV, GGOSV, and POSV) to close under spring load, 
thereby initiating the shutdown sequence previously described. 

Reset coils are provided so that the relay contacts may be returned to the 
closed "flight ready" condition for enabling subsequent attempts at normal 
engine operation. Thus, one engine malfunction does not necessarily mean 
that a restart can not be attempted. 



4-12 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Turbine over speed might be expected if propellant exhaustion occurs. 
Under this condition, there is no load on the turbopuinps, and the propel- 
lants in the start tanks are free to empty into and sustain combustion in 
the gas generator, thus accelerating the turbine wheel and turbopump 
assembly until shutdown by the overspeed circuitry of the electronic gate. 



1 



4-13 



I 



LOCKHEED MISSILES & SPACE COMPANY 



^WWMU-WfilfflW' WHIiiW,!), iiJinTi 



K!3t«W>r^?-'^T-ft»V«PVy5r*»' 



;»vrf7 J T gy?j »;yv^'iB3f'qyr,<q | y;Hffr^ ^ 



LMSC-A604141 



4.4 ROCKET ENGINE GROUND OPERATION 

The ground firing sequence is similar to that for flight operation, except 
for the use of additional ground safety features which are test facility hard- 
ware only and are removed for flight operation. The safety features (in- 
cluding necessary facility wiring and/or plumbing) are as follows: 

(1) Gas generator manifold-pressure switch (GGMPS) 

(2) Thrust-chamber pressure switch (TOPS) 

(3) Propellant tank differential-pressure switch 

(4) Emergency fuel, oxidizer, and helium.-vent valves 

All engine commands, during ground firing, are relayed through facility 
timers. If the TOPS and GGMPS do not "make" within the specified time, 
automatic termination of all electrical power to the engine is initiated at 
the facility. Should malfunction come prior to the first 143 sec of engine 
firing, the facility emergency relief ports automatically open. 

Gas Generator Manifold-Pressure Switch — The GG manifold pressure 
switch is designed to remove the facility electrical power from the engine, 
if pressure within the gas generator does not reach 312 psig (nominal) 
within 0. 5 sec after application of the start signal. This de-energizes the 
solenoid valves for shutting down the engine as described in Paragraph 4. 3. 

The GGMPS is required to prevent detonation of propellants that might have 
leaked into the GG manifold. During ground firings, when the vehicle is in 
a vertical position, the gas generator is in a horizontal position. If one of 
.the gas generator solenoid valves leak, propellant could accumulate in the 
manifold and create a potentially explosive condition. This condition does 
not exist in flight, since the liquid is free to bleed to ambient (vacuum) 
through the turbine exhaust duct. 



4-14 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Thrust-Chamber Pressure Switch — The Thrust-Chamber Pressure Switch 
initiates the engine- shutdown sequence, if thrust-chamber pressure (P^) 
does not reach 372 psig (nonninally) within two sec after application of 
engine start signal, or at any time P^ falls below 372 psig (nominal) during 
normal burn. 

Fe^cility P Pressure Switch — The facility P pressure switch actuates 
facility emergency-vent valves if at any time the fuel tank pressure drops 
within 2. 5 psid of oxidizer tank pressure with propellants in the tanks. 
This functions to preclude the possibility of excessive oxidizer tank pres- 
sure causing bulkhead reversal. 

Facility Emergency-Vent Valves — Facility emergency-vent valves are 
provided for use in the event of malfunction of the helium sphere and pro- 
pellant tanks. After first actuation of the pilot-operated helium.-control 
valve, there is no provision on the vehicle for venting the pressurization 
system. Should it become necessary to stop the engine early in the test, 
the helium sphere would continue to blow down into the propellant tanks 
and would eventually rupture them. With the addition of the facility 
emergency-vent valves, any time that the engine is shut down during the 
first 143 sec of firing, the helium sphere and propellant tanks are auto- 
matically vented. The emergency vent valves are also actuated by the 
facility P s-witch, in the event that fuel tank pressure drops within 2. 5 psid 
of oxidizer tank pressure. 

The einergency vent valves are strictly a facility item. Vehicle plumbing 
has been altered to allow access to the helium sphere pressure and pro- 
pellant tank pressure. 



4-15 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Section 5 
INSTRUMENTATION 

5. 1 INTRODUCTION 

Instrumentation requirements for the Gennini-Agena Target Vehicle (ATV) 
comprise three categories as follows: 

• Prelaunch 

• Flight teleinetry 

• In-flight status display 

Prelaunch instrunaentation monitors the status of various Propulsion System 
components to determine vehicle readiness for flight. Measurements taken 
during flight, and relayed to tracking stations by telemetry, are those which 
are used to deternnine the flight performance of the Prinnary Propulsion 
System (PPS). 

Information presented on the Agena status panel (ASP) is necessary for 
the astronaut's safety and operation of the ATV. 



5-1 



LOCKHEED MISSILES & SPACE COMPANY 



■rnvMot ■wwsTW g n ' ii i jjc iuaygw r^'^ma^funtm.v^ ■ j 



LMSC-A604141 



5. 2 PRELAUNCH INSTRUMENTATION REQUIREMENTS 

5. 2. 1 Primary Propulsion System 

Table 5-1 lists the prelaunch measuremients which are monitored while the 
vehicle is on the pad. Countdown critical transducer readings are trans- 
mitted to the control center by the vehicle umbilical J- 100 which is 
disconnected at launch. 

Launch or Hold Criteria — The first group listed in Table 5-1 concerns 
those measurements which determine if the ATV is ready for launch. 

Table 5-1 

PRIMARY PROPULSION SYSTEM PRELAUNCH 
INSTRUMENTATION REQUIREMENTS (LANDLINE) 

Transducer 
Measurement Range Maximumi Minimum 

1. Oxidizer Start "tank 

Pressure, psi to 1500 Ref. Fig. 3-5 

2. Fuel Start Tank 

Pressure, psi to 1500 Ref. Fig. 3-4 

3. Oxidizer Start Tank 

Temperature Monitor, 

OF -10 to +150 115 35 

4. Fuel Start Tank 

Temperature Monitor, 

op -10 to +150 115 35 

5. PPS Oxidizer Tank 

Temperature Monitor, 

Of 40 to 150 60 45 

6. PPS Fuel Tank 

Temperature Monitor, 

Of 40 to 150 60 45 



5-2 



LOCKHEED MISSILES & SPACE COMPANY «* 



LMSC-A604141 



Table 5-1 (Cont. ) 



Measurement 

7. PPS Oxidizer Tank 

Pressure Monitor, psig 

8. PPS Fuel Tank Pressure 

Monitor, psig 

9* Helium Gas Temperature, 

10. Helium Supply Pressure 

Monitor, psig 

11. Oxidizer Pump Lip-Seal 

Pressure Monitor, psig 



Transducer 
Range 

to 80 

to 80 

40 to 150 

to 3000 

to 25 



Maximunn 



32 



40 



165 



Minimum 



30 



38 



Ref. Fig. 5-1 



Values Derived from Above Readings Determining Launch or Hold 



1. Propellant Tank Differential 
Pressure 



Fuel over Oxidizer, 
psid 

Oxidizer over Fuel: 

Empty Propellant 
Tanks, psid 

Propellants in 
Tanks, psid 



34. 5 



to 34. 5 
5 to 34. 5 



• Additional Prelaunch Instrumentation Requirements 

1. N^ High-Pressure and Lip-Seal Valve Open Monitor 

2. Lip-Seal Valve Control (Open) 

3. Propellant Isolation Valve Control (Open) 

4. Propellant Isolation Valve (Position) 



5-3 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 

















2500 














2450 






DESIGN LOAD 


NG CURVE S 












/ 










L0> 


\DING RANGE — 


Y / 






O 2400 








/ / 






I/) 








/ / 






0. 








/ / 






at 








/ / 






Oi. 








/ 






3 








/ 






(/) 








/ 






in 








i A 






UJ 








/ / 






01 

°- 2350 








/ / 












/ / 






UJ 








/ / 






oc 








/ / 






UJ 








/ / 






X 








/ / 






□. 








f / 






t/1 








/ 






2 








/ 






3 








/ 






_l 








/ 






UJ 








/ 






X 2300 




/ 




/ 






2250 






Y 










2200 


j\^ 




/ 











40 



60 80 100 

HELIUM GAS TEMPERATURE (°F) 



120 



Fig. 5-1 Primary Propulsion System Prelaunch Helium Loading 



5-4 



LOCKHEED MISSILES & SPACE COMPANY 



!yi' ,Mj ! .tp,Wlffi^ i . 'i |,» i' g ! i» ' . ',< J ' 'f-^^i -^ ■^^ J t|) ffl g i^>i^V ' : ' -- " ^^ ''' ''--' ^ *'^""'' V*t-»t«V ? * P ^ 



ir^^«!Kp-rfl!?w??^!'r'fr-'TS*^--5'*r;?^J 



^ wt,yr^^^,^»^^^rf.'f.i9S^ 



-% 



LMSC-A604141 



Start Tanks -Start tank temperatures and pressures are used to determine 
Z:;r:^m readmess immediately prior to the vehicle launch. Transducers 
indicating start tank pressures are those which sense fuel and oxidizer venturx 
inlet pressures. Temperature-pressure relationships in the start tanks are 
established when the tanks are pressurized with nitrogen during loadmg. Any 
change in pressure, which does not relate to a corresponding change in 
temperature, may be sufficient reason for demating the Gemini -ATV 
from the booster and investigating the possibility of component leakage wxthm 
the start system. System instrumentation errors average out and do not 
affect the readings, and "Go." "No-Go" decisions are based upon the change 
in pressure only. This pressure change, with no leakage, is slight and 
influenced by temperature only. Figures 3-4a through 3-4h show fuel start 
tank and Figs. 3-5a through 3-5h depict oxidizer start tank pressure- 
temperature relationships. 

Main Propellant Tanks - Temperatures of the propellants in the main tanks 
are required to determine propellant specific gravity which varies with 
temperature. The "Go, " "No-Go" launch criteria are based on the temperature 
range of both propellants (45° to 60°F), and the temperature differential 
between the two propellants (7°F max. ). Since specific gravity of the 
propellants varies with temperature (Fig. 3-1) and the volume of propellants 
loaded is constant, the weight of impulse propellants is inversely proportional 
to temperature. Thus, if propellant temperature is too high, the mass of 
propellants loaded could be volumetrically insufficient to provide minimum 
burn time and accomplishment of mission objectives. For example: the 
weight difference between 100 ft^ of oxidizer loaded at 50°F and 100 ft 
loaded at 70°F would be approximately 125 lb, an equivalent of over 3 sec 
of main engine burn. A large temperature differential between the two 
propellants during engine firing would significantly alter the mixture ratio, 
because of density changes, and result in a lower than desired I^p. 



5-5 



LOCKHEED MISSILES & SPACE COMPANY 



.»:»»?W-?*^«V. J' «'r* "--^^ " 



LMSC-A604141 



Propellant tank pressure monitors are required for the following reasons: 

( 1) To preclude the danger of propellant tank common bulkhead 
reversal 

(2) To maintain proper liftoff ullage pressures 

(3) To preclude the possibility of exceeding the safe operating 
* pressures of the tanks 

(4) To deternnine the pressure differential between fuel and oxidizer 
tanks, which varies as shown in Table 5-1. 

Helium Sphere —Helium-sphere gas tenriperature and pressure are monitored, 
during loading, in order to accurately fill the sphere with the proper mass of 
gas (Fig. 5-1). This parameter is also required to limit pressure to a 
value connmensurate with the structural strength of the sphere. At 
temperatures above 165 F, the helium sphere structural strength decreases 
and the pressure at which the sphere will burst commensurately decreases. 
Therefore, limiting the loading rate and keeping the temperature below 
this value are necessary. 

Oxidizer -Pump Lip Seal — A pressure of 2 to 8 psig is required for lip-seal 
pressurization. This parameter is necessary to ascertain proper functioning 
of the lip pressurization system prior to launch, as a "Go, " "No-Go" 
parameter. 

Additional Prelaunch Instrumentation Requirements — To assure the 
presence of the oxidizer-pump lip-seal purge during propellant loading 
prior to launch, the position of the lip-seal valve must be known and the 
means of controlling the valve position nmust be provided. 



5-6 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Similarly, knowledge of the position and the capability of propellant 
isolation valves (PIV's) are necessary for the proper loading of the main 
propellant tanks. The PIV's must be in the open position for propellant 
loading and ascertained open immediately prior to liftoff. 

5. 2. 2 Secondary Propulsion System 

Table 5-2 lists Secondary Propulsion System (SPS) prelaunch instrumentation 
requirements and presents the qualifying range for launch. These parameters, 
when within this required range, indicate that the SPS has been loaded properly 
and should function nornnally in flight. 

Table 5-2 

SECONDARY PROPULSION SYSTEM PRELAUNCH 
INSTRUMENTATION REQUIREMENTS (LANDLINE) 











Transducer 




Measurement 




Range 


1. 


Nitrogen Gas 


Sphere (+Y), 


°F 


to 150 


2. 


Nitrogen Gas 
Tank (+Y), 


Oxidizer 
OF 




to 150 


3. 


Nitrogen Gas 
(+Y), °F 


Fuel Tank 




to 150 


4. 


Nitrogen Gas 
(-Y), °F 


Fuel Tank 




to 150 


5. 


Nitrogen Gas 
(+Y), psig 


Manifold 




to 300 


6. 


Nitrogen Gas 
(-Y), psig 


Manifold 




to 300 


7. 


Nitrogen Gas 


Sphere {+Y), 


°F 


to 150 


8. 


Nitrogen Gas 
Tank (-Y), 


Oxidizer 
OF 




to 150 


9. 


Nitro£?en Gas 


Sphere (-Y). 


°F 


to 150 



Maximum Minimum 



10. Nitrogen Gas 

11. Nitrogen Gas 



Sphere (-Y), psig to 4500 
Sphere (+Y), psig to 4500 



100 



100 



100 



100 



210 







190 



210 


190 


100 





100 





100 





See Fig. 


5-2 


See Fig. 


5-2 



5-7 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



TANK STRUCTURAL 

LIMIT ALLOWED WITH 

NO PERSONNEL EXPOSED 




60 80 

NITROGEN GAS TEMPERATURE (°F) 



Fig. 5-2 Secondary Propulsion System Prelaunch Nitrogen Loading 



5-8 



LOCKHEED MISSILES & SPACE COMPANY 



:..]\n ..> m ' mmmmi'> i - "•■" iii i . i#i . « igit!;^ii " ii' i »ii " '» u . ijiuj l ii . i ii . i ^. ";""~~ Trrr : f~ " "": ~.--— «--.».-~--~-.~>-^»^. ^-... . ...u ' .. . u-| i u. ii. ,.» .. i. ( .iu. -.!».^ 



LMSC-A604141 



5. 3 FLIGHT INSTRUMENTATION - TELEMETRY 

5. 3. 1 Primary Propulsion System Instrumentation 

Table 5-3 lists the mieasurements necessary for satisfactorily determining 
the flight performance of the Gemini-ATV Primary Propulsion System. 
Information from the transducers is relayed to ground tracking stations by 
PCM telemetry. At the stations, the data are recorded by tape recorders and 
then are played back and printed on paper tapes. Transducer locations are 
shown schematically in Fig. Z-2. 

Table 5-3 
PRIMARY PROPULSION SYSTEM FLIGHT INSTRUMENTATION 



Measurement 
No. 



Sample 

Rate 

(per sec) 



*B-1 


16 


*B-2 


16 


B-3 


16 


B-6 


96 


B-7 


16 


B-8 


16 


B-9 


16 


B-11 


16 


B-12 


16 


♦ B-139 


32 



Measurement 

Fuel Pump Inlet Pressure, psig 

Oxidizer Pump Inlet Pressure, 
psig 

Turbine Manifold Pressure 
No. 1, psig 

Combustion Chamber Pressure 
No. 1, psig 

Helium Supply Pressure, psia 

Oxidizer Tank Pressure, psig 

Fuel Tank Pressure, psig 

Oxidizer Venturi Inlet Pressure, 
psig 

Fuel Venturi Inlet Pressure, psig 

Model 8247 Engine Switch 
Group, V 



Transducer 
Range 

to 100 

to 100 

to 750 

to 750 
to 4000 
to 60 
to 60 

to 1500 
to 1500 

to 5 



* Standard Agena Measurement 



5-9 



LOCKHEED MISSILES & SPACE COMPANY 



-■tc«ffb^« » .-.-v j-<j^ i 11 Mn I ii m w m m - '^ ..^fisinranascnsQCSSf' 



■l^<3rs:'EeiW»Jt*;?31?5P*l3^^*^'^^* 



LMSC-A604141 



Table 5-3 (Cont. ) 

Sample 
Measurement Rate Transducer 

No. (per sec) Measurement R< 

■ o. 



lange 



♦B-3 1 1 Fuel Pump Inlet Temperature, F to 100 

*B-32 1 Oxidizer Pump Inlet Teinperature, 

°F to 100 

*B-3 5 DD Turbine Speed LSB''"" 

B-68 1 Helium Sphere Temperature, F to 120 

B-71 1 Oxidizer Pump Lip-Seal 

Pressure, psia to 30 

B-82 32 Fuel Valve Actuation Pressure, psig to 1500 

B-83 1 Thrust-Chamber Skin Temperature, 

Of -50 to +150 

*B-91 16 Combustion-Chamber Pressure 

No. 3, psig 475 to 550 

1 Fuel Tank Temperature No. 1, °F -15 to +170 

1 Fuel Tank Temperature No. 2, °F -15 to +170 

32 Propellant Isolation Valves 

Open/Closed, v to 5 

3 2 Turbine Manifold Pressure No. 2, 

psia to 120 

1 Oxidizer Tank Skin Temperature 

(+Z), OF -15 to +170 

1 Oxidizer Tank Skin Temperature 

(-Z), °F -15 to +170 

1 Oxidizer Start TarLk Temperature, 

Op -10 to +150 

1 Fuel Start Tank Temperature, °F -10 to +150 

32 Oxidizer Injector Pressure, psig to 1000 

1 Nozzle Extension Skin Temperature 

No. 1, °F -200 to +800 

B-185 1 Nozzle Extension Skin Temperature 

No. 2, Op -200 to +800 



B- 


■96 


B- 


•97 


B- 


-130 


B- 


■132 


B- 


-136 


B- 


-137 


B- 


-141 


B- 


-142 


B- 


-148 


B- 


-184 



* Standard Agena Measurement 
=*L,east significant bit 



5-10 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Note from Table 5-3 that the Gemini-ATV instrumentation requirements 
are much more extensive than those for the standard Agena D (standard 
Agena D measurements prefixed with an asterisk). The reasons for these 
differences are that the Model 8247 is a new engine without previous flight 
history, and present mission profiles require that the ATV have an extended 
orbital period. Thus, more diagnostic measurements are necessary in 
o^rder to ascertain the behavior of the propulsion system during the orbital 
period, the pre-dock period, orbital maneuvers, and the post-dock period. 

Pump Inlet Pressures — Fuel and oxidizer pump inlet pressure transducers 
(B-1 and B-2, respectively) are basic Agena instrumentation needed for deter- 
mining total propellant flow rate which is necessary to calculate Specific 
Impulse (I ). These transducers serve as backup instrumentation to tank 
top pressures (B-8 and B-9) but may not be used as a substitute for them. 
B-1 and B-2 measurements give total average inlet pressures (suction 
pressures) for the turbopumps during an engine burn, which is one of the 
factors in determining total flow rate. Pump inlet pressures are necessary 
in making calculations for the Subsystem Technical Analysis Report (STAR), 
which evaluates the propulsion system performance in flight. 

Turbine Manifold Pressure No. 1 — Turbine manifold pressure No. 1 (B-3) 
is used to assess operation of the gas generator, turbine manifold, and 
turbine wheel. 

Combustion-Chamber Pressure No. 1 -Combustion chamber pressure 
No. 1 (B-6) is used in determining thrust-chamber start and shutdown 
transients. This measurement provides information necessary for 
determining the start and shutdown impulse for each engine burn. The 
values obtained from B-6 are used only during engine start or shutdown; 
measurement B-91 (expanded scale) is used for steady-state thrust-chamber 
performance. 



5-11 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Heli um Supply Pressure — The helium supply pressure (B-7) is monitored 
during first burn only; this measurement indicates proper operation 
of the pyrotechnically operated helium-control valve (POHCV). The 
pressure blowdown, as shown by B-7, will be connpared to blow-down 
predictions run by computer at LMSC, Sunnyvale. This comparison will 
show total flight-pressurization system performance and indicate any 
abnormal areas. The ineasurement will also be utilized to determine any 
leakage of the high-pressure helium system during launch and prior to 
first burn. 

This is the same transducer which is monitored on landline instrumentation 
during ground fill of the helium sphere. 

Propellant Tank Pressures — Oxidizer and fuel ullage pressure measure- 
ments {B-8 and B-9, respectively) provide information on a number of 
parameters which are important to the success of the mission. The 
blow-down curves, generated by B-8 and B-9, will be assessed and 
compared to predictions (run during computer studies at LMSC, Sunnyvale) 
to ascertain that proper pressurization system operation has taken place 
during the blow-down phase of pressurization. During orbital flight, the 
B-8 and B-9 measurements will be monitored to determine any leakage in 
the propellant system upstream of the propellant isolation valves. Prior 
to any main engine firing, B-8 and B-9 will be checked to assure that 
sufficient pressure is present in the propellant tanks, to preclude the 
possibility of cavitation, upon start of the turbopumps. This parameter 
will be especially scrutinized prior to any main engine operation by the 
astronaut when he is in the vicinity of or docked to the Gemini- AT V, 
If leakage in the fuel system allows the oxidizer tank pressure to become 
greater than the fuel tank pressure, a bulkhead reversal could possibly 
take place resulting in damage and possible vehicle destruction. Monitoring, 
by ground stations, of B-8 and B-9 will yield information necessary to the 
astronaut's safety. 



5-12 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



» 



Propellant Venturi Inlet Pressure - Venturi inlet pressure transducers 
(B-11 and B-12) are basic instrumentation required for computing gas 
generator (GG) flow rates. The GG flow rates are needed to compute the 
total engine flow rate which is used in determining engine performance. 
These measurements will be the only means of determining pressure on 
the liquid side of the start tanks during the orbital coast period. Informa- 
tion from these transducers will be used to ascertain if the start system 
has been properly recharged during a previous engine burn. When an 
astronaut is docked to the vehicle, prior to his primary propulsion system 
firing initiation, this measurement must be determined to be within limits 
for a safe PPS start. These are the same transducers which are monitored 
on landline for start tank status prior to liftoff. 



h 



Model 8247 Engine Switch Group - The Model 8247 switch group (B- 139), 
more accurately called engine "tell-tales, " provides the only means of 
determining whether or not a turbine overspeed relay trip has occurred. 
This switch group has four voltage levels (Fig. 5-3) indicating the operational 
status of the relay contact matrix contained within the engine -mounted 
electronic gate. Prior to engine operation, the voltage level is zero. 
Upon application of electrical power to the engine, with normal engine 
operation, the level rises to 0. 94v (nominal) and remains there throughout 
normal engine operation. In the event of turbine overspeed, relays K-1 
and K-2 trip and the voltage level rises to 4. 7v (nominal). Should one of 
the relays fail to open on overspeed shutdown, or fail to close on reset, 
this would be indicated by voltage levels for relay position Nos. 1 and 2. 
Relay position No. 1, at a nominal voltage level of 1. 95, indicates that 
relay K2 is open and K^ is closed. If K2 is closed and Kj is open, this will 
be indicated by a voltage level of 3.2 (nominal) as the second relay position. 



5-13 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



u 



o 
> 



OVERSPEED 4.7 VOLTS 



RELAY POSITION NO. 2 
(3.2 VOLTS) 



RELAY POSITION NO. 1 
(1.95 VOLTS) 



0.94 VOLTS 



NORMAL ENGINE OPERATION 



ENGINE START SIGNAL 

THESE VALUES ARE BASED ON 28V DC ON THE VEHICLE 
BUSS AND WILL VARY AND BE CALIBRATED FROM VEHICLE 
TO VEHICLE. 



Fig. 5-3 Engine Switch Group Nominal Voltage Levels 



5-14 



LMSC-A604141 



Propellant-Pump Inlet Temperature — Fuel and oxidizer pump inlet 
temperatures (B-3 1 and B-3Z, respectively) are standard Agena-D 
instrumentation necessary for determining specific gravity of the propellants, 
which, in turn, is used in calculating propellant flow rate. Monitoring 
these parameters during coast periods following an engine burn will 
indicate the effect of heat transfer from the gearbox and turbine assembly. 
This is a diagnostic -type measurement to help determine the effects of 
PIV's and the venting of propellants overboard following an engine burn. 

Turbine Speed — The turbine speed mieasurement (B-35) is basic instrumen- 
tation necessary for determining total fuel and oxidizer flow rates which 
are needed to determine engine performance. This parameter would 
indicate actual turbine speed in the event of a turbine overspeed condition. 
For diagnostic reasons, this measurement will show the start and shutdown 
transients. These data may be compared with PTVA data generated at 
Santa Cruz Test Base or Research and Development data generated at BAG 
to ascertain normal performance. 

Helium Sphere Temperature — The helium sphere temperature measurement 
(B-68) is required for calculating the sphere polytropic blow-down exponent 

N '. Since "N" is used in predicting the blow-down characteristics of the 
sphere, this measurement is necessary for determining flight pressure- 
tinne histories which are compared with those assunned for the preflight 
calculations. This measuremient is nnonitored on landline prior to launch. 

Oxidizer -Pumip Lip-Seal Pressure — This nr^easurement (B-71) is required 
to assure that the required 2. to 8. psig oxidizer lip-seal pressure is 
maintained any time propellants are present in the engine. 



5-15 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Fuel Valve Actuation Pressure — The fuel valve actuation pressure (B-82) 
will indicate when the fuel valve opens during start transients. This meas- 
urment is required due to the Model 8247 engine starting sequence (fuel 
lead) being different than that of the Model 8096 engine. To determine 
nominal engine performance, this parameter will be reviewed. 

Thrust-Chamber Skin Teinperature — This measurement (B-83) is 
required for diagnostic reasons to assure that thrust-chamber skin 
temperature does not exceed the maximum permissible. 

Combustion-Chamber Pressure No. 3 — The suppressed zero cumbustion- 
chamber pressure No. 3 {B-91) is basic instrumentation which is necessary 
for calculating total engine thrust. This is one of the required inputs for 
the STAR. 

Fuel Tank Skin Temperatures — The two main fuel tank skin-temperature 
pickups (B-96 and-B-97) are located on the +Z and -Z axes, respectively. 
These naeasuremients are required to show the effect of exposure to space 
temperature and/or exposure to sun on vehicle propellants and to determine 
the vehicle absorptivity/emmissivity coefficient (a/f). These measurements 
are Gemini-peculiar due to the extended orbital period. 

Propellant Isolation Valves Open/Closed — The position of the propellant 
isolation valves is indicated by measurement B- 130, The PIV's will be 
opened 4 sec prior to, and closed 4 sec after, PPS burn. This parameter 
is diagnostic and required to determine proper valve operation. 

Turbine Manifold Pressure No. 2 — This measurement (B-132) is diagnostic 
and is used to determine proper turbine and gas generator operation. 



5-16 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Oxidizer Tank Skin Temperature - The two main oxidizer tank skin- 
temperature transducers (B-136 and B-137) are mounted on the +Z and 
-Z axes, respectively. These measurements are needed to show the 
effect of exposure to space temperatures and/or exposure to sun on vehicle 
propellants and to determine the vehicle absorptivity/emmissivity coefficient 
(a/f). These measurements are Gemini-peculiar due to the extended 
orbital period. 

Propellant Start-Tank Temperatures -The oxidizer and fuel start tank 
temperatures (B-141 and B-142, respectively) are required to determine the 
mass of gas present within the start tank, and together with venturi inlet 
pressures will indicate whether leakage has taken place in either the gas or 
liquid side of the start tanks. These are redline parameters, required 
prior to liftoff , and are monitored on landline. 

Oxidizer Injector Pressure - This parameter (B- 148) will be used to 
determine opening time of the main oxidizer valve. The measurement 
will also reflect proper turbopump/thrust-chamber operation. 

Nozzle-Extension Skin Temperature - The two nozzle-extension skin temper- 
atures pickups (B-184 and B-185) are located on the +Z and the -Z axes, 
respectively. These measurements are needed to determine if the maximum 
permissible temperatures of the material (titanium) of the nozzle extension 
has been exceeded. 

5. 3. Z Secondary Propulsion System 

Table 5-4 lists the measurements necessary for satisfactorily determining 
the flight performance of the Gemini-ATV Secondary Propulsion System. 
Information from the transducers is relayed to ground tracking stations by 
PCM telemetry. Transducer locations are shown schematically in Fig. 4-3, 



5-17 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Measurement 
No. 


Sample 

Rate 

(per sec) 


B-200 


16 


B-201 


16 


B-202 


16 


B-203 


16 


B-204 


16 


B-205 


16 


B-208 


16 


B-209 


16 


B-212 


32 


B-213 


32 


B-214 


32 


B-215 


32 


B-216 


1 


B-223 


16 


B-224 


16 


B-225 


16 


B-233 


1 


B-234 


1 


*2. V = 
3.0 V = 


-Y on 
+Y on 


**2. V = 
3.0 V = 


Unit II on 
Unit I on 



Table 5-4 

SECONDARY PROPULSION SYSTEM 
FLIGHT INSTRUMENTATION 



Measurement 



Gas Sphere Pressure (+Y), psia 

Gas Sphere Pressure (-Y), psia 

Tanlc Manifold Pressure (+Y), psia 

Tank Manifold Pressure (-Y), psia 

Oxidizer Feed Pressure (+Y), psia 

Oxidizer Feed Pressure (-Y), psia 

Fuel Feed Pressure (+Y), psia 

Fuel Feed Pressure (-Y), psia 

Unit I Chamber Pressure (+Y), psia 

Unit I Chamber Pressure (-Y), psia 

Unit II Chamber Pressure (+Y), psia 

Unit II Chamber Pressure (-Y), psia 

Nitrogen Gas Sphere Temperature 
(+Y), °F 

Start Valve On/Off (+Y & -Y), v 

Units I and II Pilot Solenoid On/Off 
(+Y), V 

Units I and II, Pilot Solenoid On/Off 
(-Y), V 

Unit I Bipropellant Valve Tempera- 
ture (+Y), °F 

Unit I Bipropellant Valve Tempera- 
ture (-Y), °F 



Transducer 
Range 

to 4500 

to 4500 

to 300 

to 300 

to 300 

to 300 

to 300 

to 300 

to 100 

to 100 

to 120 

to 120 

-50 to +150 
to 5* 

to 5** 

to 5** 

-50 to +250 

-50 to +250 



5-18 



LOCKHEED MISSILES & SPACE COMPANY 



fT:lt,',>"^j! !*?!!"''■■!'' ^ ■•■ '. ' 



|f^^P(KPJEP"^7FT^|?'^T'^W*^^T'*^'^ 



LMSC-A604141 



Measurement 
No. 

B-235 

B-236 

B-240 
B-241 
B-246 
B-247 
B-248 
B-249 
B-253 

B-260 

B-261 

B-262 

B-263 

B-266 

B-267 

B-268 

B-269 



Sample 

Rate 

(per sec) 



o. 



Table 5-4 (Cont, ) 



M easurement 

Unit II Bipropellant Valve Tempera- 
ture (+Y), °F 

Unit II Bipropellant Valve Tempera- 
ture (-Y), OF 

Unit I Injector Temperature {+Y), F 

Unit I Injector Temperature (-Y), 

Unit I Skin Temperature (+Y), °F 

Unit I Skin Temperature (-Y), F 

Unit II Skin Temperature (+Y), F 

Unit II Skin Temperature (-Y), F 

Nitrogen Gas Sphere Temperature 
(-Y), °F 

Unit I Oxidizer Feed- Line 
Temperature (+Y), °F 

Unit II Oxidizer Feed-Line 
Temperature (+Y), °F 

Unit I Oxidizer Feed- Line 
Temperature (-Y), F 

Unit II Oxidizer Feed -Line 
Temperature {+Y), °F 

Unit I Fuel Feed-Line^ 
Temperature (+Y), F 

Unit II Fuel Feed-Line 
Temperature (+Y), °F 

Unit I Fuel Feed-Line^ 
Temperature (-Y), F 

Unit II Fuel Feed-Line 
Temperature (-Y), °F 



Transducer 
Rang e 

-50 to +250 

-50 to +250 
- 100 to +200 
-100 to +200 
32 to 2800 
32 to 2800 
32 to 2300 
32 to 2300 

-50 to +150 

-50 to +150 

-50 to +150 

-50 to +150 

-50 to +150 

-50 to +150 

-50 to +150 

-50 to +150 

-50 to +150 



5-19 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Secondary Propulsion System instrumentation is discussed in this report, 
because of the relationship of SPS to proper operation of the PPS. 

Nitrogen Sphere Pressure —The +Y and -Y modules transducers (B-200 and 
B-201) are required to determine the mass of gas remaining during orbit 
for evaluating system capability. These transducers are also utilized to 
analyze and predict the condition of the Agena status panel (ASP), "SPS Hi' 
anS "SPS Lo" lights. Nitrogen sphere pressure will be monitored continually 
for indications of high-pressure gas-system leakage. 

Tank Manifold Pressure — The +Y and -Y modules measurements {B-20Z and 
B-203) are required to verify energization of the start valve and gas pressure- 
regulator operation as follows: 

(1) Provides verification that tardus are pressurized to operating 
pressure prior to thrust-chamber operation 

(2) Analyzes condition of ASP lights 

(3) Provides "second source" information for performance calculations 
(propellant flow rates) 

(4) Is monitored continually during orbit to detect indications of 
low-pressure gas-systenn leakage. 

Oxidizer Feed Pressure — The +Y and -Y modules transducers {B-204 and 
B-205) are required to calculate oxidizer flow rates, which are a function 
of pressure drop across the thrust-chamber assembly. 

Fuel Feed Pressure — The +Y and -Y modules measurements {B-208 and 
B-209) are required to calculate fuel flow rates, which are a function of 
pressure drop across the thrust-chamber assennbly. 



5-20 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC A604141 



P 



I 
1 



Unit I Chamber Pressure -The +Y and -Y modules transducers (B-212 and 
B-213) provide the best single source of information for evaluating Unit I 
thrust-chamber assembly operation. Chamber pressures are required to • 
calculate Unit I fuel and oxidizer flow rates, which are a function of pressure 
drop across the thrust chaiT^ber assembly. In addition, these measurements 
are required to calculate thrust and specific impulse. 

TTnU TI Chamber Pressure - The +Y and -Y modules transducers (B-214 and 
B-215) provide the best single source information for evaluating Unit II 
thrust-chamber assembly operation. Chamber pressures are required to 
calculate Unit II fuel and oxidizer flow rates, which are a function of 
pressure drop across the thrust chamber assembly. In addition, these 
measurements are required to calculate thrust and specific impulse. 

Nitroeen Sphere Temperature - The +Y and -Y modules measurements (B-216 
and B-253) are required to determine mass of gas remaining during orbit 
for evaluating system capability. Information on SPS module orbital history 
is also provided by these measurements. 

Start Valve ON/OFF (+Y and -Y) - Verification of proper system sequencing - 
i.e., "Ready" command -is provided by this parameter {B-223) which has a 
range of to 5 volts. Applicable voltages are 2. 0, confirming that the -Y 
start valve has been commanded to open, and 3. ^indicating that the +Y 
start valve has been commanded to open. 

Units I and II Pilot Solenoid ON/OFF Module (+Y) - Verification of proper 
system sequencing -i.e., "Fire" command - is provided by this parameter 
(B-224) which has a range of to 5 volts. Applicable voltages are 2. 0, 
confirming that the Unit II solenoid has been commanded to energize, and 
3. Vindicating that the Unit I solenoid has been commanded to energize. 



5-21 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Units I and II Pilot Solenoid ON/OFF Module (-Y) - Verification of proper 
system sequencing -i.e., "Fire" command - is provided by this paran^eter 
{B-225) which has a range of to 5 volts. Applicable voltages are 2.0, con- 
firming that the Unit II solenoid has been commanded to energize, and 3. 0, 
indicating that the Unit I solenoid has been commanded to energize. 

Unit I Bipropellant Valve Temperature - The +Y and -Y modules measure- 
ments (B-233 and B-234) are required to provide data ( 1) on probability of 
safe starts -i.e.. propellant is not frozen, (2) for determining coast period 
requirements, and (3) on orbital thermal history of the SPS module. 

Unit I Injector Temperature - The +Y and -Y modules measurements (B-240 
and B-241) are required to provide data (l) on probability of safe starts at 
low temperatures (below 0°F) and (2) on orbital thermal history of the SPS 
module. 

Unit I Chamber Skin Temperature - The +Y and -Y modules measurements 
(B-246 and B-247) provide data on the normality of Unit I thrust chamber 
operation. 

Unit II Chamber Skin Temperature - The +Y and -Y modules measurements 
(B-248 and B-249) provide data on the normality of Unit I thrust chamber 
operation. 

Unit I Oxidizer Feedline Temperature - The +Y and -Y modules measure- 
ments (B-260 and B-262) are required (1) to obtain specific gravity factors 
in Unit I oxidizer flow-rate determination. (2) to provide data on orbital 
thermal history of SPS modules, and (3) for use to predict performance 
degradation of Unit I thrust-chamber assembly, caused by oxidizer cavita- 
tion in the thrust-chamber trim orifices. 



5-22 



LOCKHEED MISSILES & SPACE COMPANY 



-*f |jj ii; a ;i ^ ' g|MjkJ i |«g ! y ( g e y,?^ 



,'j y j |"!' ^tf' ^ q^^;^'WT^ f ^ 1 ^ . J 'T i »^_ ^ !fi^ ^ -^ •*' 



LMSC-A604141 



Unit II Oxidizer Feedline Temperature — The +Y and -Y nnodule measurements 
{B-261 and B-263) are required (1) to obtain specific gravity factors in 
Unit II oxidizer flow-rate determination and (2) to provide data on orbital 
thermal history of the SPS modules. 

Unit I Fuel Feedline Temperature — The +Y and the -Y modules measure- 
ments (B-266 and B-268) are required (1) to obtain specific gravity factors 
in'Unit I fuel flow-rate determination and (2) to provide data on orbital 
thermal history of the SPS nnodules. 

Unit II Fuel Feedline Temperature — The +Y and the -Y modules measure- 
ments {B-267 and B-269) are required (1) to obtain specific gravity factors 
in Unit I fuel flow-rate determination, and (2) to provide data on oribital 
thernnal history of the SPS modules. 



5-23 



LOCKHEED MISSILES & SPACE COMPANY 



,y ! My. i iyB i ^, ' t ' Lji^f.-j j s yggr*^ 



?^^.5pigpfljljj^j5ji(n^-Tn«i7;^^ 



*^!«»iwi^»5E«W^!BrrJi^^^;^. 



LMSC-A604141 



5.4 AGENA STATUS DISPLAY PANEL 

The Gemini-ATV Status Panel (ASP), shown by Fig. 5-4, is mounted on the 
forward end of the Target Docking Adapter of the Agena Target Vehicle where 
it is visible to the astronauts in the Gemini spacecraft during and after the 
docking maneuver. The panel displays information on the status and safety 
of the Agena propulsion, guidance, electrical power, and docking systems. 
Originally, only eight Agena parameters were to be displayed in the Gemini 
spacecraft; however, the number of parameters increased to the point that the 
spacecraft no longer had the space or weight capability to accomodate 
them. Accordingly, the panel was placed on the Target Docking Adapter. 



The ASP system consists of a display panel (Fig. 5-4) with nine display 
lights and three analog dials and the necessary circuitry which is distributed 
throughout the Gemini-ATV. When not in use, this system is normally 
de-energized in order to save power; however, the PPS and SPS Time 
Remaining Clocks are energized whenever the PPS or SPS engines fire. 

Three of the twelve parameters displayed on the ASP panel, indicate PPS 
status and three indicate SPS status. The Primary Propulsion System displays 
are as follows: 

• PPS Burn Time Remaining Clock 

• "MAIN" Red Light 

• "MAIN" Green Light 

The Secondary Propulsion System displays are as follows: 

• SPS Burn Time Remaining Clock 

• "SEC HI" Green Light 

• "SEC LO" Green Light 



5-24 



LOCKHEED MISSILES & SPACE COMPANY 



y)^M.!BVi->yijj|jpig jT gj^j,j) l ^y^^ 



re w- ' »'-rv^j"»M»-jjW'U ." tJ4.i : J . V ' f7 ^w> 'J jAiww*'i!a i iWiV ./^ ? ^! » »s ^ #^^ 



PSi^jr?sB;<r«3—'«>,"iv^ 



LMSC-A604141 




0^ 

>^ 

nj 
1—1 

P. 
tn 

D 

CO 

-t-> 

4-> 

to 

A) 
PI 
<D 
M 

< 

I 



bJD 



5-25 



LOCKHEED MISSILES & SPACE COMPANY 



w«^ ' ^M W U ■ Jl !flwgr ^ '3;^^^^ff^^l••-■ ^ aiKS^P »■W f ^^».^^;T*^ . ';^ >^ ^ ^ ^ , 



LMSC-A604141 



Requirements received from the astronauts stated that a dimming circuit 
must be added to the ASP system, because the panel lights are found to be too 
bright when the spacecraft docks with the ATV on the night side of the earth. 
Consequently, a dimming circuit has been added to reduce or increase the 
brightness to the desired level. The dimming circuit functions on all lights 
except MAIN Red which is always bright when "ON" for reasons outlined 
below. In order to improve reliability, two lamps are incorporated in each 
indicator light. 

5. 4. 1 Time Remaining Clocks 

Primary Propulsion System Clock - The Primary Propulsion System burn-time 
remaining display (Fig. 5-4) is a dc motor-driven clock with the short hand 
indicating the number of minutes (up to six) and the long hand the number of 
seconds (up to 60) of PPS burn-time remaining. The amount of time which the 
Primary Propulsion System will burn with the amount of fuel and oxidizer 
loaded have been previously calculated and the clock preset to this time via an 
AGE reset capability provided in the clock. This clock, with a decreasing 
time indication, runs whenever the PPS is burning and continually provides 
an indication of the amount of burn time left for operation of the system. The 
Primary Propulsion System clock is started by the "PPS Thrust Initiate" 
command, and stopped by the "PPS Thrust Cutoff" command. 

Secondary Propulsion System Clock — The Secondary Propulsion System 
burn-time remaining display is a dc motor -driven clock similar to the one 
used for the PPS time display. The SPS clock can be run at two different 
speeds depending upon whether the Unit II high-thrust units (200 lb force) or 
the Unit I low-thrust units (16 lb force) are burning. The ratio used to differ- 
entiate between a Unit I and a Unit II firing is 12 to 1. For every second of 
Unit II firing, the clock registers 1 second less of time remaining, and for 
every 1 second of Unit I firing, the clock registers 1/12 of a second less of 
time remaining. This ratio of 12 to 1 was determined by taking a ratio of 



5-26 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



the total mass -flow rate of propellants to the Unit II thrust chambers and 
dividing by the total mass -flow rate of propellants to the Unit I thrust chambers, 
The SPS Unit II thrust-time remaining is presented directly, whereas the clock 
reading must be multiplied by 12 to find the Unit I thrust-time remaining. The 
available equivalent burn-time for the SPS Unit II thrust cham.bers is preset 
into the clock via an AGE reset capability provided in the clock. 

Originally the burn-time remaining analog displays were to be pulse-driven 
for accuracy; however, since these displays were a 5 to 10 percent indication, 
the high accuracy of a pulse-driven system was not necessary. Therefore, 
NASA directed that dc -driven clocks be used. 

5.4.2 MAIN (RED) LIGHT 

Light Operation - The MAIN (red) light (Fig. 5-5) may be turned "ON" by one 
of three ways: Turbine overspeed, hydraulic high-pressure-switch actuation, 
and fuel-oxidizer tank delta-pressure switch actuation, 

Propellant Tank AP Switch —A differential pressure switch, mounted in the 
forward section of the vehicle and used to sense both fuel and oxidizer 
tank-top pressures, is set at 3 ±2 psid fuel-tank over oxidizer -tank. If the 
fuel tank pressure is less than 3 ±2 psid above the oxidizer tank, the only 
remaining prerequisite to light operation is the Gemini-ATV unregulated 
28 vdc bus energization. Prior to launch and propellant tank prepressurizing, 
a blanket pressure (5 psi or less) is maintained on both propellant tanks 
equally; therefore, any time that the 28 vdc unregulated bus is on, the MAIN 
(red) light energizes. There is no cause for concern, since the light is 
extinguished when propellant tanks are pressurized. 

Sufficient pressurization gas is loaded aboard the 3653 -in. p sphere to meet 
all mission requirements and maintain tank pressure differentials above the 
actuation value of the 4P switch. Therefore, if the AP switch causes 
the MAIN (red) light to come on, the system has a leak. 

5-27 



LOCKHEED MISSILES & SPACE COMPANY 

■ ii i i «ji; i ftj) i g ,ijj ppy.» )|i ti v vrfc i i.fri t tiHV#-; ' wv, »i ji ^^ 





LMSC-A604141 



m 






ill 



^1% 



*+- 



J 



•ri 



61 






>— |ss HI 



^ 



^ 



V- 




^ 



' S^ ^ b 5 



UJSg 



^ 






I— 1 

0> 


P! 


ri 


u 


rt 


a) 


0* 


w 


t>N 


13 


(ti 


f3 




a 


•iH 

Q 


(1) 




■<-> 


w 


M 


Hi 


!>^ 


4J 


W 


4-> 


C 


W 


o 

•H 


a 


w 


rii 


■— 1 


4-J 

'0 


Ph 


!>^ 


u 


rn 


In 




Oh 


d 




o 


>^ 


• r-l 


u 


UJ 


(ti 


-t 


H 


P^ 


• r-4 


O 


!-< 


u 


Ph^ 


in 




in 




W) 




■ iH 




h 





5-28 



'^RWSPRp^jap^*! 



•^'■i.'^'?!^_w^^vi^ t^r^^. 



LOCKHEED MISSILES & SPACE COMPANY 



^^^^^^^If^f^j^^^i^f^^^f^^^^'!'^^^-^:' 



LMSC-A604141 



A study of propellant time-pressure history was initiated and nominal curves, 
together with upper and lower 3-sigma limits, were developed (Fig. 6-3). 
Further studies were made with the aim of obtaining tnaximiim warning time 
in the event of a pressurization system malfunction (leakage). An optimum 
oxidizer isolation valve closing-time was determined at 318 sec (Fig. 6-3) 
after initiation of the Model 8247 engine start signal. This time (318 sec) 
provides maximum capability for sensing imminent propellant tank bulkhead 
reversal (a possible catastrophic failure). The minimum pressure differ- 
ential between tanks, 9 psid, occurs toward the last few seconds of PPS burn 
capability (using the lower 3-sigma fuel tank pressure and the upper 3-sigma 
oxidizer tank pressure). Utilizing the upper limit of the 4P switch setting, 
the amount of fuel tank pressure leakage prior to 4P switch actuation is 
4 psi. At the opposite extreme, utilizing the upper 3-sigma limit of fuel tank 
pressure and the lower 3-sigma limit of the oxidizer tank pressure and taking 
the minimum allowed 4P switch actuation point, a pressure leakage of 24 psi 
can occur before a switch actuation takes place. Therefore, if the 4P switch 
actuates, a fuel tank pressure leakage of 4 to 24 psi will have occurred. 

Hydraulic High-Pressure Switch -Referring to Fig. 5-5, note that when 
either command 201 or 211 is given, 28 vdc is supplied to normally open K-16 
relay contacts. Once the "PPS Thrust Initiate" command is given, a ground 
return is supplied to relay coil K-13 contacts to the coil of relay K-l6, 
Relay K-16 does not energize unless a ground path is created through the "OR" 
gate by either turbine overspeed or by the hydraulic high-pressure switch 
dropout. A 1. 5 sec (measured from "PPS Thrust Initiate" command) delay 
is incorporated to allow time for the hydraulic pressure switch to build up 
to 1500 ±20 psi without igniting the MAIN (red) light. If this pressure switch 
fails to pickup during the allowed 1. 5 sec or drops below switch setting any- 
time during normal engine operation, the circuit immediately supplies the 
required ground, through the "OR" gate, to relay K-16, closing K-16 
normally open contacts, and thereby lighting MAIN (red) on the ASP. 



5-29 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Nominal hydraulic-discharge pressure is Z810±60psig. The hydraulic pres- m 

sure switch setting of 1500 psi was picked as a mean between engine control- m 

lability and indication of gross hydraulic systenn malfunction. Selection of the 
1. 5 sec delay to build up hydraulic -actuation pressure was based upon data ' M 

generated during hydraulically controUed-engine testing indicating an approxi- 
mate 1.0 sec delay maximum, over the specification extremes of temperature, 
to reach 1500-psi hydraulic -discharge pressure. 

Turbine In Overspeed — Sinailar to the hydraulic high-pressure switch, 

the turbine in over -speed signal (Fig. 5-5) can trigger the "OR" gate and allow M 

relay coil K-16 to energize by supplying a ground return at the "OR" gate, thus •* 

lighting the MAIN (red) light on the ASP. Nominal turbine speed is 24, 800 rpm. ^ 

The signal to the "OR" gate is supplied by means of a frequency-sensitive relay ^ 

systenn which triggers at 27, 000 rpm. This trigger value was obtained by 

analyzing engine test data and establishing a 3-sigma level below which a H 

family of Model 8247 rocket-engine start transients would lie. If the turbine 

continues to accelerate, an automatic turbine overspeed shutdown is initiated M 

by the engine -mounted electronic gate (see Paragraph 2. 2. 1 of this report) at 

29, 500 ±500 rpm. m 

To determine rise times of the turbopum^i speed, a study was initiated with «| 

calculations based on assunaed pump unloading factors. The critical shaft speed ^ 

is 31, 000 rpm (no dwell allowed), and the rotor rotational stress lower-limit is 

41, 000 rpm. The fastest rise time (approximately 40 ms) is calculated fronn M 

27, 000 (ASP "MAIN" red "ON") to 29, 500 rpm (automatic cutoff) and occurs 

with conapletely dry turbopumps. The slowest rise tinne (approxinnately m 

500 ms) is calculated from 27, 000 to 29, 500 rpm and occurs during a ixial- 

function in which the fuel valve closes and the oxidizer continues to unload M 

overboard. A most likely condition to cause turbine overspeed would be where * 

maximunn utilization of all PPS propellants was attemipted and propellant m 

exhaustion occurred, with oxidizer exhausting first and fuel exhausting soon gj 

afterward; a 12 -lb fuel bias is included during propellant tank loading. 

(See Section 3 of this report). In the aforementioned case, the turbine speed H 

5-30 1 



LOCKHEED MISSILES & SPACE COMPANY " 



LMSC-A604141 



rise time is from 27, 000 to 29, 500 rpm in approximately 180 ms. Present 
Gemini-ATV mission profiles do not purposely include the intent to maximize 
usage of all propellants. 

Reaction To MAIN Red During Flight -If the MAIN (red) light comes "ON" 
when the Prinnary Propulsion System is not firing, the malfunction indicated 
would be due to 4P switch dropout and would indicate a possible tank bulkhead 
reversal. Should the light come "ON" during nornnal PPS burn, immediate 
renaedial action is required by the spacecraft. Shutdown of the PPS must be 
initiated by issuing spacecraft command 500. If the MAIN (red) light goes 
"OFF" after command 500, the next ground station should have information, 
via telennetry, as to the cause of malfunction; i.e., whether the malfunction 
was due to turbine over speed or hydraulic system malfunction. Should the 
light remain "ON", possible tank bulkhead reversal may take place, and the 
astronaut must take imnnediate appropriate action until the next ground 
tracking station can access the probable cause. Paragraph 5.2 of this report 
describes instrumentation available to determine possible failure modes. 

Although the ASP has two light-intensity levels available, dim {command 211) 
or bright (command 201), the MAIN (red) light always illuminates bright when 
a malfunction occurs. 

5.4.3 MAIN (Green) Light 

Light Operation - The MAIN (green) light (Fig. 5-5) may be turned "OFF" 
by one of three ways: fuel tank pressure -switch dropout, oxidizer tank 
pressure-switch dropout, and hydraulic low-pressure-switch dropout. 
Figure 5-5 shows that either command 211 or 201 can supply power to the 
MAIN circuit and illuminate the MAIN (green) light, if the hydraulic pressure 
in the reservoir is above 50 psi, the oxidizer tank-top is above 15 psia, and 
the fuel tank -top is above 15 psia. 



5-31 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



H ydraulic Low-Pressure Switch — The hydraulic low-pressure switch, set at 
50 ±5 psia, is utilized to assure that proper hydraulic residual-pressure 
remains in the hydraulic reservoir. If this pressure should be lost, due to 
a leak or other reasons, the hydraulic pressure would not build up properly 
when the "PPS Thrust Initiate" command is sent. This could lead to in- 
stability of thrust direction during firing, 

0>»idizer Tank Pressure Switch — The oxidizer tank pressure switch, set at 
1 5 ±2 psia, is utilized to inform the astronaut if oxidizer tank pressure is 
decaying, due to a leak or other reasons. The turbopump requires a nninimum 
head for proper operation and for suppressing cavitation. Oxidizer pump 
cavitation takes place below 10 psia at a propellant temperature of 60°F. 

Fuel Tank Pressure Switch — The fuel tank pressure switch, set at 15 ±2 psia, 
is utilized to inform the astronaut if fuel tank pressure is decaying, due to 
a leak or other reasons. The turbopunnp requires a minimum head for proper 
operation and for suppressing cavitation. Fuel pump cavitation takes place 
below 13 psia at a propellant temperature of 60 °F. 

This switch is actually redundant with the AP switch located in the MAIN (red) 
circuit. If the MAIN (green) light is "ON", indicating sufficient oxidizer tank 
pressure to suppress cavitation; and the MAIN (red) light is not "ON", 
indicating the fuel tank pressure is at least 3 ±2 psi above the oxidizer tank, 
proper operation without the possibility of cavitation is assured. 

Reaction to MAIN Green During Flight — The MAIN (green) circuit is designed 

to convey maximum information to the astronaut concerning the status of the 

Genaini-ATV and the pernaissibility of the Primary Propulsion System to fire. 

If the light is not "ON", no PPS firing should be attempted until the next 

ground station can, via telemetry, deternnine the cause. If the light should 
g£ "OFF" during PPS burn, cavitation of the turbopump may be imminent, 

and the engine may either be manually shut down, or the astronaut may rely 

on the engine -mounted turbine overspeed device to perform, the shutdown. 

5-32 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



5.4.4 SEC LO (GREEN) LIGHT 

Light Operation - The SEC LO (green) light may be turned "OFF" by one or . 
naore of four ways: +Y nitrogen-gas tank low-pressure -switch dropout, 
-Y nitrogen-gas tank low-pressure-switch dropout, +Y manifold pressure- 
switch dropout, and -Y manifold pressure-switch dropout. Figure 5-5 shows 
that either command 211 or 201 will supply power to the SEC LO circuit and 
illuminate the SEC LO (green) light, if either the +Y or -Y nitrogen spheres 
are above 360 psia, or either the +Y or -Y manifolds are above 170 psia. 

Manifold Pressure Switch — The nomi. 1 pressure experienced in the +Y or 
-Y manifold (identical modules) is 204 psia (nominal) regulated pressure. 
Allowed regulator deviations reduce this pressure to 197 psia (nominal), and 
the 3 -Sigma product variations further reduce the m.inimuna manifold pressure 
levels to 190 psia (nominal). A value of 170 ±5 psia (dropout) for switch set- 
ting was finally calculated to allow 20 psi for temperature variations of the 
gas in the manifold while in orbit. 

Nitrogen Gas Tank Low-Pressure Switches (+Y and -Y) — The +Y and -Y 
nitrogen spheres are each nominally loaded to 4000 psig. The low-pressure 
switches must operate in the aforementioned system environment but actuate 
(dropout) at 360 ±20 psia. Low-pressure switch setting is calculated on the 
basis that the minimunn. pressure would sustain one additional 150 sec 
(full duration) continuous Unit I operation. Due to the low propellant-flow 
rate, the final sphere pressure is 210 psia, allowing for proper regulator 
operation and, therefore, nominal thrust ratings throughout the firing duration. 

Reaction To SEC LO (Green)Light During Flight - The SEC LO (green) light 
should be used in conjunction with the SPS clock as an indication of thrust-time 
available for Unit I thrust chambers. If the SEC LO (green) light goes "OFF" 
during an SPS Unit I normal burn sequence, the next ground station should be 



5-33 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



consulted for information concerning remaining burn-time available for 
Unit I operation. Should the SEC LO (green) light g£ "OFF" when Secondary 
Propulsion System firing is not taking place, the loss of the light may be 
indicative of sphere or manifold leakage, and the next grotmd station should 
be consulted for the cause of light loss. 

If the SEC LO (green) light is "OFF" and the SPS clock indicates time 
remaining in excess of one full-duration Unit I firing, a manifold leak may 
have developed. The nianifold is the system most susceptible to leakage, 
since a small volume of gas is trapped between the start solenoid valve and 
tank bellows (Fig. 4-7). If a leak has developed, and the SEC LO (green) 
light goes "OFF" due to the pressure switch dropping below 170 psi, a start 
valve open command may be sent, and if SEC LO relights, a Unit I firing may 
be initiated. 

5.4.5 SEC HI (Green) Light 

Light Operation — The SEC HI (green) light may be turned "OFF" by one or 
more of four ways: +Y nitrogen-gas tank high-pressure-switch dropout, 
-Y nitrogen-gas tank high-pressure-switch dropout, +Y nnanifold pressure- 
switch dropout, and -Y manifold pressure-switch dropout. Figure 5-5 shows 
that either command Zll or 201 will supply power to the SEC HI circuit and 
illuminate the SEC HI (green) light, if either the +Y or -Y nitrogen spheres 
are above 1110 psia, or either the +Y or -Y nnanifolds are above the 170 psia. 

The +Y and -Y manifold pressure switches are discussed in Paragraph 5.4.4, 
and as indicated by Fig. 5-5, these switches operate identically for the 
SEC LO and SEC HI light circuits. 

Nitrogen Gas Tank High-Pressure Switches (+Y and -Y) - The +Y and -Y 
nitrogen spheres are each nominally loaded to 4000 psig. The high-pressure 
switches must operate in the aforementioned systena environment but 



5-34 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



actuate (dropout) at 1110 ±20 psia. High-pressure switch setting is calculated 
on the basis that the mininraum pressure would sustain one additional 50 sec 
(full duration) continuous Unit II operation. For proper SPS regulator 
operation during a Unit II burn, the final sphere pressure must not drop 
below 400 psia for the relatively high propellant-flow rate. 

Reaction To SEC HI (Green) Light During Flight - The SEC HI (green) light 
should be used in conjunction with the SPS clock as an indication of thrust- 
time available for Unit II thrust chambers. If the SEC HI (green) light goes 
"OFF" during an SPS Unit II normal burn sequence, the next ground station 
should be consulted for information concerning remaining burn -time available 
for Unit II operation. Should the SEC HI (green) light go "OFF" when 
Secondary Propulsion System firing is not taking place, the loss of the light 
may be indicative of sphere or manifold leakage, and the next ground station 
should be consulted for the cause of light loss. 

If both SPS LO and SPS HI (green) lights go "OFF" simultaneously, this 
would be indicative of either +Y or -Y manifold leakage (manifold pressure- 
switch dropout). Should both lights go "OFF" and the Secondary Propulsion 
System is firing, then the SPS firing should be immediately terminated and 
the next ground station consulted for data indicating the cause. If both lights 
go "OFF" and the Secondary Propulsion System is not firing, then the SPS 
start valve open command (command 56l) may be sent to open the start 
solenoid and repressurize the manifold to a value above 170 psi (manifold 
pressure-switch pickup pressure). Should the lights (HI and LO) remain 
"ON", an SPS firing may be attempted. 



i,| ii »|i]i !i . iiip .»«yj»a»ij miMwu ) ... I. . 11 i j fjujfj i . | .Ml^^i* ' -W ' " " :)"^' )^ " ?-^ ^'y'?r? ?'y ' 'w-^5 !' ! ; »« <^q°^^''''^*'"?^'''P^'^*^^"'' -" ""' •'■'■''■v-'- 



5-35 



LOCKHEED MISSILES & SPACE COMPANY 

,..^.,,,,,,,,,,,,^,,,5,^^.p„.:,^ 



LMSC-A604141 



5. 5 PRESSURE SWITCHES 

Pressure switches are mounted on various PPS and SPS components to 
warn the astronaut when undesireable conditions develop within the 
propulsion systems. These switches operate at predetermined pressures 
and either extinguish a green light or light the MAIN (red) light, while 
maintaining a display to indicate the current status of the propulsion 
systenn. 

5. 5. 1 Fuel Over Oxidizer 4P 

The fuel over oxidizer 4P switch is a normally closed differential pressure- 
sensing switch mounted integrally on the main propellant tank assembly. 
This switch has high- and low-pressure inlets connected to the fuel and 
oxidizer tanks, respectively. The higher pressure on the fuel side holds 
the switch open. If fuel tank pressure drops to within 3 ±2 psid of 
oxidizer tank pressures, the switch returns to its normally closed position. 

The switch employs twin mechanical bourdon tubes which are independant 
mechanically from each other. A design prerequisite was imposed upon 
the switch, such that any single malfunction would not allow a common 
passage between fuel and oxidizer systems. The switch has characteristics 
as shown in Table 5-5. 

5. 5. 2 Fuel Tank Pressure Switch 

The fuel tank pressure switch is a normally open pressure-actuated switch, 
connected to the ullage section of the fuel tank. This switch is connected in 
series electrically with the oxidizer tank pressure switch and the hydraulic 
low-pressure switch. Design characteristics are presented in Table 5-5. 



I 



5-36 



LOCKHEED MISSILES & SPACE COMPANY 



w3«pj?5ij;pp??(P^-'*^- 



LMSC-A604141 



I 
i 



PRESSURE SW 



Table 5-5 
.ITCHES DESIGN CHARACTERISTICS 



5-37 



Pressure 



150 
200 
150 
700 



|^aelTH^^.Oxidi5erJ:a^^^ ^ ^^ ,00 

Pressure, Range, psid 
Proof Pressure (elements), psid 
Burst Pressure (elennents), psid 
Line Pressure (nnaximuni), psig 
Burst Pressure (case), psig 
Fueljraixk_Pre£SureJv^^ ^ ^^ ^^q 

Pressure Range, psia ^^^ 

Proof Pressure (elenrient), psia ^^^ 

Burst Pressure (elenaent), psia ^^ ^^ 

Switch Point, psia 

g2d^^:e^J^r^^^l:^Sl}iI^^^^^^^^ to 100 

Pressure Range, psia ^^^ 

Proof Pressure, psia ^^^ 

Burst Pressure, psia ^^ ^^ 
Switch Point, psia 

HYdrauli£i:iOW^P££lSHI^^^ to 100 

Pressure Range, psia 
Proof Pressure (element), psia 
Burst Pressure (element), psia 
Burst Pressure (case), psia 

SPSJiitra^£ILS2ll£££J^^^ to 4000 

Pressure Range, psia 
Burst Pressure (element), psia 
Burst Pressure (case), psia 



IBO 
200 
700 



8000 
4000 



LOCK 



HEED MISSILES 8c SPACE COMPANY 



LM3C-A604141 



Table 5-5 (Cont. ) 



SPS Nitrogen Sphere High-Pressure Switch 
Pressure Range, psia 
Burst Pressure (element), psia 
Burst Pressure (case), psia 

SPS Manifold Pressure Switch 
Pressure Range, psia 
Burst Pressure (elennent), psia 
Proof Pressure (element), psia 
Burst Pressure (case), psia 



Pressure 

to 4000 

10,000 

4000 

to 300 
600 
450 
700 



5. 5. 3 Oxidizer Tank Pressure Switch 

The oxidizer tank pressure switch is a normally open pressure -actuated 
switch, connected to the ullage pressure of the oxidizer tank. This switch 
is connected in series electrically with the fuel tank pressure switch and 
the hydraulic low-pressure switch. Design characteristics are shown in 
Table 5-5. 

5.5.4 Hydraulic Low-Pressure Switch 

The hydraulic low-pressure switch is a single-pole double-throw pressure- 
actuated switch set at 50 ±5 psia. Electrically, it is connected in series 
with the oxidizer and fuel tank pressure switches. Design characteristics 
are presented in Table 5-5. 



5-3i 



LOCKHEED MISSILES & SPACE COMPANY 



^^ ^ i%m^af90 ^ '» m ^,\i. f WMl^^'HI^^^^^ 



' ? ' ' ; f» » i ;y ?*» * y «* yy -g^ ra'?w''ffy »H?.*i!-i^8« ' j ' UV- a! H--'- ' -'r*'!T g p g a r!e ^ ?; i«W *' ''fttJ'fW ^ ■ 'i ^.. ' ! '-■', ^-7*'y?! 



LMSC-A604141 



5. 5.5 Secondary Propulsion Systenn Nitrogen Sphere Low-Pressure Switch 

The SPS nitrogen sphere low-pressure switch is a single-pole single-throw 
pressure-actuated switch set at 360 ±20 psia. Design characteristics are 
shown in Table 5-5. 

5. 5.6 Secondary Propulsion System Nitrogen Sphere High-Pressure Switch 

The SPS nitrogen sphere high-pressure switch is a single-pole single-throw 
pressure-actuated switch set at 1110 ±20 psia. Design characteristics are 
shown in Table 5-5. 

5. 5. 7 Secondary Propulsion System Manifold Pressure Switch 

The SPS manifold pressure switch is a single-pole double-throw pressure- 
actuated switch set at 170 ±5 psia. Design characteristics are shown in 
Table 5-5. 



5-39 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Section 6 

PRIMARY PROPULSION SYSTEM PERFORMANCE 
AND DESIGN CHARACTERISTICS 



6. 1 INTRODUCTION 

This section discusses the propulsion subsystem performance capability and 
the parameters governing engine performance and power level in vacuum 
flight. The method for predicting flight parameters and for determining 
flight performance are also presented. 

6. 2 ENGINE REQUIREMENTS 

The pressurization and propellant feed systems supply propellants to the 
engine pump inlets' at a sufficient pressure to permit the propellant pumps to 
operate above the minimum permissible pressure required to suppress 
cavitation. Hence, the minimum permissible pump inlet pressures of the 
engine are the key values in establishing the design of the pressurization 
system and propellant tanks. The pump inducers attached directly to the pump 
impellers (Figs. 2-19, 2-20, and 2-21) have been incorporated to reduce 
the required inlet pressure. In effect, the inducers act as a boost pump, 
adding head to the propellants to suppress cavitation in the main impeller. 
The engine pump minimum total inlet pressures are shown in Figs. 6-1 
and 6-2. At nominal turbine speed and at a propellant temperatures of 60 F, 
the minimum permissible pump inlet pressures are 13.05 psia for fuel and 
9. 80 for the oxidizer. 



6-1 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



FUEL; UDMH 
PER WIL-D-25604B-1 
ATMOSPHERIC PRESS: PSIA 



q;22 



£20 

fc 
^18 



?)6 



RANGE OF ROTATIONAL SPE EDS 
N) • DOS OF DESIGN SPEED 
Nj - DESIGN SPEED 
N3.90'?OF DESIGN SPEED 













1 












/ 1 












/ / 1 










/ 


v/ 










/} 


^ / 








^ 


/ / 


^ 


N) =27 927 


CPU 


^^^ 


h ^ 


V 








^^ 




/ 




"Nj =25,339 


RPM 


^ 


-^ 






N3 =22,8^9 


^Ea--- 


"^ 





















60 80 100 120 

PROPELLANT TEMPERATURE-- 'F 



UO 



Fig. 6-1 Fuel Pump Minimum Inlet Pressure 
Versus Propellant Temperature 



OXIDIZER: IRFNA 
PERMIL-P-725'IE-l 
ATMOSPHERIC PRESS: PSIA 



RANGE OF ROTATIONAL SPEEDS 
N] . 110% OF DESIGN SPEED 
N2- DESIGN SPEED 
N j- 90' !; OF DESI GN SPE ED 



24 
22 












/ 












7/ 












/// 


18 
16 
14 












// 












/ 










y/ 








^ 


^ ^^ 






12 

10 




^^^^^"^^ 


"^^A 






Nj = 14 4 RPU ^-^ 










— |ijjJ2,9 
























6 
4 



























80 100 

prCpellant temperature^ 'F 



PROPI 



140 



Fig. 6-Z Oxidizer Pump Minimum Inlet Pressure 
Versus Propellant Tennperature 



6-2 



LOCKHEED MISSILES & SPACE COMPANY 



^..yJiy l lil4^ l jjjL■ | ^ ^! jjBy^ l ;^t;'';fi^ij,* T J l A ' ^ 



:-^,- ^ '" i^ )i ijj. ff ;.l f ^, ^ .. ' i? yiy'7 ;, T i«yj S j ] i.: ■■> t f ww ii y -i!^<wty?yy<y.:y 



lR' i: w»r i i y. Ty^'??1g!«yijlipgg' 



LMSC-A604141 



6. 3 ORIFICE-FED PRESSURIZATION SYSTEM 

The primary tool used for predicting orifice -fed pressurization system 
(OFPS) performance, and in part, the Gemini ATV primary propulsion 
system flight performance, is the computer program originally set up to 
determine feasibility. The capability of adequately predicting tank and 
sphere pressure-time histories was denaonstrated at the Santa Cruz Test 
Base by PTV static firing tests. 

The actual process involved in sphere blowdown is a comprise between 
adiabatic and isothermal. To compute these blowdown curves properly, 
a polytropic exponent, varying with time, was utilized. 

A sufficient bias is established between the fuel and oxidizer tank (Fig. 6-3) 
to preclude common bulkhead reversal. 

Examination of the starting transients (Fig. 6-3) indicates the following: 
The oxidizer and fuel tank pressures decrease initially because ullage 
volumes increase prior to actuation of the POHCV. At approximately 
2 seconds after the control valve opens, the tanks repressurize to their 
maximum value and then decay gradually during first burn blowdown. 
After first burn shutdown, both tank pressures rise again. At 3 18 seconds 
after engine start signal, the oxidizer tank is isolated fronn the helium 
sphere. The fuel tank pressure continues to rise while oxidizer tank 
pressure remains constant; thus increasing fuel over oxidizer pressure 
differential, providing a greater safety margin against bulkhead reversal. 
During subsequent burns the tank pressures continue to blowdown until the 
mission requirements are completed. Tank 3-sigma variations indicate the 
bands in which the propellant tank pressures will remain during operation. 



6-3 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




(VISdTs3drt5s3ad1(NVliNVn3ldOiJir 



i § 



(visd) dbnssaad BddHds MiinaH 



6-4 



LOCKHEED MISSILES & SPACE COMPANY 



« 




s 




•H 




H 




0) 




M 




::* 




to 




to 




<u 




M 




Ph 




Ai 




C 




oJ 




H 




■!-> 




d 




r— I 


T! 


t-H 


0) 


(LI 


4-i 


PL, rt 


O 


h 


l^ 


<!) 


0^ 


a, 


T3 


o 




•1-1 




'I 


(U 


rr 


^^ 


rn 


(U 




4i 


1 


a, 




ro 


UJ 




lU 


s 


.1-1 


:s 





•i-i 


+-> 


I— 1 


rn 


(i> 


.rH 


XX 


CO 




vD 




W) 




• t-i 




h 





P ..W4'*:j'J ' g g?pgW WWII > % i J.il., l l ^tWW.W>4P!fWJ^Ma^^^^ 



LMSC-A604141 



6. 4 SUBSYSTEM PERFORMANCE IN VACUUM FLIGHT 
6. 4. 1 General 

The propulsion subsystem perfornnance in vacuunn flight is defined as the 
performance of the engine integrated into the propulsion system and sub- 
jected to the varying pump inlet conditions which result from tank blowdown 
*and from operation in the accelerating vehicle. Hence, vehicle weight is 
a parameter affecting engine performance and, therefore, the numerical 
performance valu'es corresponding to subsystem performance will vary 
for individual flights. This difference results from the action of tank 
blowdown and acceleration on the pump inlet pressures, and consequently, 
on flow rates, mixture ratio, and thrust. The numerical values for the 
variation of these parameters are a function of vehicle weight and other 
operating conditions. For predicting the nominal flight performance of 
any specific vehicle, the engine performance, as obtained from the engine 
acceptance test data, is extrapolated to flight conditions using the engine- 
governing equations given in Paragraph 6. 8. The parameters of thrust, 
propellant consumption, and mixture ratio are machine -computed as a 
function of time by an iterative process feeding back the effects of vehicle 
acceleration until propellant exhaustion occurs. To illustrate this 
procedure, the steps for calculating a sample point are outlined in the 
description which follows. 

6. 4. 2 Determining Flight Performance Parameters 

To determine the flight performance parameters of thrust, propellant con- 
sumption, and mixture ratio the operating conditions at the pump inlets 
and the turbine speed, as a function of these operating conditions, must be 
established. 



6-5 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



The pump inlet operating conditions affecting the performance paranneters 
of the engine are propellant density and pump inlet pressure. The effect 
of temperature, other than on density, has been empirically determined as 
negligible. For computing the nominal flight performance, propellants of 
nominal composition, and at a teniperature of 60 F, are assumed. 

The instantaneous punnp inlet pressures for both punrips at any time are 
obtained from the following basic equation: 



P. T , = P, , - AP,. + P 

inlet tank lines ace 



where; 



tank 



AP 



lines 



ace 



= the tank top pressure at the instant considered 

= the sum of the liquid and pneumatic line losses for 

fuel or oxidizer 
= the acceleration pressure head of the propellants. 



The line drops are design constants independent of the other variables. The 

acceleration pressure (P ) is a function of the liquid head (ft) above the 

ace , 

pump inlet, the propellant density (lb/ft ), and the vehicle acceleration 
(g's). It is related to these parameters by the following equation: 



ace 



aha 
144g 



where: 



A = propellant density (lb/ft ) 

h = head (ft) 

a = instantaneous acceleration 

g = local gravitational constant (varies with altitude). 



6-6 



LOCKHEED MISSILES & SPACE COMPANY 



wji!Wytfey;M'' w j<fiwu'».t iiw < iii> i i ) i iy | ij|uj»t)^^ 



= ,;y > »J ,i:;p ,, iiii t.j , i l iliny ' . : %gti»y!S !F;''J^ 



T«»*^?T7?l^ws?^?^S!39^»5!;iW7?^T^>»'^!;0^ 



LMSC-A604141 



To determine the instantaneous value of h, the volume of propellants in the 
tank at the specified time and the relationship between the head and tank 
volume must be established. The volume of propellants in the tank is 
obtained by subtracting the volume of propellant consunned from the volume 
initially loaded in the tanks. The head-versus -volume relationship is a 
function of the geometry of the tank configuration. Typical curves for the 
integral tank are presented in Fig. 6-4. 

The instantaneous acceleration is equal to the quotient of the thrust divided 
by the vehicle weight. As these parameters are interdependent, the value 
of acceleration nnust be deternnined by an iteration process. The procedure 
is to estinnate a value for thrust and calculate the corresponding accelera- 
tion: 



a = 



W 
v 



where: 

W = vehicle weight 

V 

This value of acceleration is connbined with the instantaneous head and 

propellant density to determiine the acceleration pressure (P ). The 

ace 

acceleration pressure is substituted in the basic equation to determiine the 
pumip inlet pressure. 

After the pump inlet pressures and propellant densities are determined, 
the turbine speed can be determined fronn the engine -governing equations 
given in Paragraph 6. 8. The flow rates and thrust are then deternnined 
from the relationships also given in Paragraph 6. 8. 

This value of thrust is used to deternnine a new value for acceleration, and 
the calculation is iterated until the assumed thrust checks the derived 
thrust. Then, based on the computed flow rates and assumed tinne interval, 
succeeding points are calculated until propellant exhaustion or, in the case 
of a comimand shutdown, when the desired velocity increment is attained. 

6-7 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



12 



10 



H 
U. 



< 

tu 

X 









y 






FUEL TANK 




y^ 








^ 






y 


1 


OX 


IDIZER 7 


"ANK-v 


^ 


y 




y< 




^ 








/ 














FUEL 




OXIDIZER 




20 40 60 80 100 

TANK VOLUME, FT.^ 

Fig. 6-4 Liquid Head Versus Volume for Propellant Tanks 



6. 4, 3 Determination of Performance Capabilities 

The criterion selected for determining performance capability is stage velocity 
increment as a function of payload. The velocity incrennent capability is 
based on minimum specific impulse. The payload capability is defined as 
the burnout weight less the weight of all of the propulsion system components 
and the propellant tank assembly. The entire calculation is based on the 
rocket equation for velocity increment assuming horizontal flight: 



W 

V = I e In =Tr TTTT- 

sp ^ W - W 



where: 



W = burnout weight of vehicle 
o ° 

W = weight of propulsion systenn and propellant tanks 
P 

6-8 



LOCKHEED MISSILES & SPACE COMPANY 



'->i i., ' <m^ ii :ftt^t^,itf\.m ' iiif' ' t^Mj f ^i 'f' ^ ' 'm\i-y j.u.. i - i .. « 



'^^^^''pT^IWWMr'^PJ^ 



LMSC-A60414I 



6. 5 ENGINE VACUUM PERFORMANCE 

The Subsystem B specification values for the performance parameter of 
thrust, burn time, specific impulse, and mixture ratio are presented in 
Table 6-1. 

The predictability values shown in the table are the 3-sigma variations 
of vacuum performance for the primary engine parameters. 

Table 6-1 
ROCKET ENGINE PERFORMANCE AT ALTITUDE 



Engine 

Performance 

Parameter 


Rated 


Limits of 
Set Perfornr^ance 


Predictability 


Guaranteed 
Vacuum 
Performance 
Limits 


Min, 


Max. 


Min. 


Max. 


Thrust, lb* 


16, 000 


-2.5% 
15, 600 


+2. 5% 
16,400 


±3. 08% 
±493 


-5. 58% 
15, 107 


+5. 58% 
16,893 


Specific 

Impulse, 

sec'"''" 




291.5 




±0. 15% 
±1.5 


290.0 




Mixture 

Ratio, W /W^ 
o' f 


2.57 


-1.5% 
2. 53 


+ 1. 5% 
2. 61 


±0. 6% 
±0.015 


-2. 1% 
2. 515 


+2. 1% 
2. 625 


Burn Time, 
sec 


240.0 













*Figs. 6-5 and 6-6 present thrust variations with altitude. 

**Fig. 6-7 presents I change with altitude. 

sp 



6-9 



LOCKHEED MISSILES & SPACE COMPANY 




LMSC-A604141 



16.000 



NO THRUST CHAMBER NOZZLE FLOW SEPARATION 
MIXTURE RATIO - 2.57 ±1.5% 



15,000 



14,000 



Q 

§13,000 
o 

Q. 



g 12,000 

X 



11,000 



10,000 



9,000 





/ 


^ 




v 

\--16,0 


00 






/ 


























































1 
































ESI IMAT c 

FLOW SEF 


D REGION U 
ARATION (1£ 


,875 FT) 



40 



80 120 160 200 

ALTITUDE • THOUSANDS OF FEET 



240 



Fig. 6-5 Variation of Rated Thrust with Altitude 



6-10 




LOCKHEED MISSILES & SPACE COMPANY 



'SBSIww^aiiwifo^fOfjjfiiw^iwww^i^^ 



«aw»aB;p(?PB^5rrMp«Wt-!;f^^lw7!f»^ 



LMSC-A604141 



Z 

U 
a: 



X 



Q 

ai 
< 

Z 

o 



110 
105 
100 

95 

90 

85 

80 

75 

70 

65 

60 ; 




ESTIMATED REGION OF NOZZLE 
FLOW SEPARATION (18,875) 



40 



80 120 160 

ALTITUDE (1000 FT) 



200 



240 



A^ — I 



NOTE: 



CURVES INCLUDE ONLY VARIATIONS DUE TO 
PROPELLANT DENSITY LIMITS, MANU- 
FACTURING TOLERANCES, AND SPEED 
CONTROL AND NO THRUST-CHAMBER NOZZLE- 
FLOW SEPARATION. 



LEGEND 

MAXIMUM AT +10°F 
MAXIMUM AT +32°F 
MAXIMUM AT +60°F 
RATED AT H60°F 
MINIMUM AT +60°F 
MINIMUM AT+100°F WITH„ 

PROPELLANTS AT +90 F' 
MINIMUM AT +140°F 



Fig. 6-6 Estimated Rocket Engine Thrust 

Limits Over Operating Requireinents 



6-11 



LOCKHEED MISSILES & SPACE COMPANY 



^Sm!«<* <*■*"'"<*■**■ '^^' 



LMSC-A604141 



300 




120 160 

ALTITUDE (1000 FT) 



NOTES: 



NO THRUST-CHAMBER NOZZLE FLOW SEPARATION 
MIXTURE RATION = 2.57 ±1.5 PERCENT 



Fig. 6-7 Variation of Rocket Engine Minimum 
Overall Specific Impulse with Altitude 

6-12 



LOCKHEED MISSILES & SPACE COMPANY 



:r^?^fW5J?!SSf'i?KP^»'PBW»^"*^ 



LMSC-A604141 



6. 6 NOMINAL PERFORMANCE DATA 

The engine model specification requires only the iTiiniinum acceptable value 
for specific impulse. Hence, the nominal specific impulse and the nominal 
flow rate for the engine are not specified values. The noininal value of 
specific impulse is an empirically derived parameter based on the average 
value obtained by a statistical evaluation of the data fronn a series of engine 
firings. As specific impulse is not a ineasurable parameter, it must be 
indirectly obtained by evaluating the parameters of thrust chamber charac- 
teristic velocity, c*, and nozzle thrust coefficient, C^: 



sp 



g 



These parameters are obtained by measurements of thrust, flow rates, 
throat area, nozzle exit area, and chamber pressure as follows; 



C''- = 



c t^ 



F + P A 

^F p~r 

c t 

The average or nominal values for c* and C^ are 5, 220 ft/sec and 1. 829 
respectively. 

The minimum acceptable specific impulse for the XLR-81-BA 13 engine is 
291. 5 seconds. As engines currently being delivered are meeting this 
requirement, this figure may be considered nominal. 

The nominal flow rate for the engine is defined as the rate which will 
result in nominal rated thrust at nominal specific impulse: 



W. 



sp 
6-13 



LOCKHEED MISSILES & SPACE COMPANY 



•.■■—^^rjvr^^^r vmmhmi ja gfyf^, '-'■"'. 



^ .^> W^yj.i»iH »' tf^* 'V°y*^^''^^ vn?i]R»9l»-.^ 



' ■'J. ^^ V'VM ; g;WKl^ ,>^l i! '?:T?;'! J !>;J'' » y j« ra ^ '^^M^J^ ''j ^^^ 



LMSC-A604141 



6.7 EXHAUST DUCT THRUST 

The altitude value of the exhaust duct thrust can be computed as the sum 
of the gas momentum and the pressure force acting at the duct exit: 



ex 



(P A ) + 
ex ex 




For sonic flow at the duct exit, the exhaust duct thrust will be 200 lb. 

6. 8 ROCKET ENGINE PERFORMANCE PREDICTION (PRE-FLIGHT) 

In order to predict engine performance accurately, Gemini ATV variations 
from the standard Agena D must be taken into consideration. These 
variations are caused by: 

• Nonstandard turbine exhaust duct length and shape 

• Nonstandard hydraulic motor (gimbal) flow rates 

• Nonstandard sphere size utilized in vehicle pressurization 
system^ 

• Variation of propellant temperature. 

A computer program combines these variations with previous flight 
experience and Bell Aerosystems Company (BAC) engine acceptance test 
data to predict flight performance. 

Rocket engine performance at altitude can be calculated based on an 
average of the sea level corrected data presented in the BAC engine 
acceptance log book. Rocket engine performance at altitude is defined by 
the following averaged paranneters: 



6-14 



LOCKHEED MISSILES & SPACE COMPANY 



"*^i'iJ','|",1.!W,"M*;.!fliaw)«H'« '' rr 



^.^r'F',.,.>yg-yif>^»i .;i i L is. y . 






LMSC-A604141 



Mixture ratio at altitude 



R 



• Total mass flow rate at altitude 

• Thrust at altitude 

• Specific impulse at altitude 

• Turbine speed at altitude 



W 



00 



00 



spoo 



N„ 



CO 



Each of these param^eters miust be corrected from sea level to vacuum 
conditions with the exception of I , which is assumed at a nonainal Z91. 5. 

SP -;- — 

The thrust is derived analytically; i. e. , using W , I , and turbine exhaust 

■^ ■' ° 03 spoo 

thrust. Figures 6-8, 6-9, 6-10, and 6-11 show pump design characteristics. 



6. 8. 1 Methods of Prediction 



6. 8. 1. 1 N -- Turbine Speed at Altitude 



N = (K ) (N 



n 



si) 



where: 



N, 



K 



n 



N 



si 



= turbine speed at altitude 

= computer constant 

= turbine speed at sea level corrected. This value is 
available fromi BAG Engine Log Book. 



6. 8. 1. 2 Wg3 ->- Mass Flow Rate at Altitude 



"^ w si 



6-15 




LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



+60°F 



PROPELLANT: IRFNA 
PER MIL-P-7254E-1 
TEMPERATURE: 
SPECIFIC GRAVITY: 1.570 
VAPOR PRESSURE: 2.1 PSIA 
VISCOSITY: 32 X 10'* 

SLUG/FT SEC 
ATMOSPHERIC PRESSURE: PSIA 



RANGE OF ROTATIONAL SPEEDS 

N^ . 110% OF DESIGN SPEED 

N2 - DESIGN SPEED 

N3 - 90% OF DESIGN SPEED 

DESIGN FLOW: 39.12 LB/SEC FOR THRUST CHAMBER 

0.205 LB/SEC FOR GAS 
GENERATOR 



1.200 




10 



20 30 40 

TOTAL SUCTION PRESSURE - Pjo - PSIA 



Fig. 6-8 Oxidizer Pump Flow Versus 
Total Suction Pressure 



e 



6-16 



LOCKHEED MISSILES & SPACE COMPANY 



'^"y^ fyj ' T^ ■rr «/-.y.'" ' *i» fr'^?'rrTy ^-"^yy?? ? >,' ^*^ ;-' 4 'r ' ;'ygy?^ 'T»T'.TBr-»;g»ffTO^pisw>ji<)^«)B:!?w!pE^^ 



LMSC-A60414] 



1.200 



PROPELLANT: UDMH 
PER M1L-D-25604B-1 
TEMPERATURE: +60°F 
SPECIFIC GRAVITY: 0.795 
VAPOR PRESSURE: 1.9 PSIA 
VISCOSITY: 11.9 x lO"^ 

SLUG/FT SEC 
ATMOSPHERIC PRESSURE: 



PSIA 



RANGE OF ROTATIONAL SPEEDS 

N, . 110% OF DESIGN SPEED 

Nj - DESIGN SPEED 

N- .90% OF DESIGN SPEED 

DESIGN FLOW: 13.90 LB/SEC FOR 
THRUST CHAMBER 
1.370 LB/SEC FOR GAS 
GENERATOR 

0.806 LB/SEC FOR HYDRAULIC 
MOTOR (RETURN TO SUCTION 
LINE) 



1.100 



o 



z 

UJ 



'^I.OOO 



0. 

3 



.900 



.800 




TOTAL SUCTION PRESSURE • Pjp • PSIA 



Fig. 6-9 Fuel Pump Flow Versus Total Suction Pressure 



6-17 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



1600 



1500 



1400 



1300 



^ 1200 
I 1100 

CO 

UJ 
Qi 

a. 

U 1000 



Q- 900 

s 

CL 

800 



700 



600 



PROPELLANT: IRFNA 
PER MIL-P7254E-1 
TEMPERATURE: +60"f 
SPECIFIC GRAVITY: 1.570 
SUCTION PRESSURE AS NOTED 
VAPOR PRESSURE: 2.1 PSIA 
VISCOSITY: 32 x lO"'^ 

SLUG/FT SEC 
ATMOSPHERIC PRESSURE: 



O PSIA 



RANGE OF ROTATIONAL SPEEDS 
N^ " 110% OF DESIGN SPEED 
Nj ° DESIGN SPEED 
N3 = 90% OF DESIGN SPEED 



500 











Pjo = 28.6 PSIA 
N, = 15,851 RPM 


"\ 




\ 


Pgo = 24.0 PSIA ^*>.,.^^ 
N2 "= 14,410 RPM 


\ 






\ 




PSo"^ '9.8 PSIA "--v. 

N3 = 12,969 ^v^^^ 


tut 


) 



















10 



20 30 40 

FLOW TO THRUST CHAMBER - LB/SEC 



50 



Fig. 6-10 Oxidizer Pump Static Pressure Rise 
Versus Flow Rate at Constant Speeds 



6-18 



LOCKHEED MISSILES & SPACE COMPANY 



?^»*^^l■^¥y'^' ^ ^ »t^l^l8^^i^^ f i JHyiuia i feM kt u^piliiwiyg^tj,. I B 



rr aw .'; ^ J 'ti j , 'i»]»R -'Ba;i^ vJ .jjtj )ij^ i uffM » .w^^ 



LMSC-A604141 



PROPELLANT: UDMH 
PER MIL-D-25604B-1 
TEMPERATURE: -HSO^F 
SPECIFIC GRAVITY: 0.795 
SUCTION PRESSURE AS NOTED 
VAPOR PRESSURE: 1.9 PSiA 
VISCOSITY: 11.9 X 10"* 

SLUG/FT SEC 
ATMOSPHERIC PRESSURE: PStA 



1500 



RANGE OF ROTATIONAL SPEEDS 

N, » 110% OF DESIGN SPEED 

Nj " DESIGN SPEED 

N3 " 90% OF DESIGN SPEED 




9 19 n 12 13 

FLOW TO THRUST CHAMBER • LB/SEC 



14 



15 



Fig. 6-11 Fuel Pump Static Pressure Rise Versus 
Flow Rate at Constant Speeds 



6-19 



LOCKHEED MISSILES & SPACE COMPANY 



,»K«Mi«:31^Jf.i«»;. : . 



LMSC-A604141 



where: 



W, 



00 



K 



w 



W 



si 



= mass flow rate at altitude 

= computer constant 

= mass flow rate at sea level corrected. This value is 
available from BAG Engine Log Book. 



6. 8. 1. 3 R^ — Mixture Ratio at Altitude 



Ra. = (R ,) (K ) 
SI r 



where: 



R 



00 



R 



si 



K 



= mixture ratio at altitude 

= average sea level corrected mixture ratio. This value 

is. available in the BAG Engine Log Book, 
= computer constant 



6. 8. 1.4 F^ ~ Thrust at Altitude 



CD 



(I ) (Woo) + F ^ 
spno ^ ' exh 



where: 



thrust at altitude 



Sp RO 



exh 



specific impulse at altitude. This is a fixed value 
Ibf 
Ibm/sec 



of 291. 5 



= exhaust duct thrust (constant ZOO lb) 



6-20 




LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



6. 8. Z Example of Predicted Flight Values 

Values taken from BAG Engine Log Book: 

I =291.5 ^-r^ 

sp Ibm/sec 



N 



^ = 24, 497 rpm 



W , =53. 34 
si 

R , =2. 55 
si 



Values taken from conaputer study: 



K = 1.00208 

n 

K. = 1.06500 
w 

K = 1.02100 

r 



Solve for N„ 



N 



CD 



N, 



N, 



(K ) (N -) 
n si 

(1.00208) (24,497) 
24, 548 rpm 



Solve for W 



00 



W^ = (1.06500) (53.34) 
W„ = 56.81 



6-21 



LOCKHEED MISSILES & SPACE COMPANY 



Byg«j y . » j!iyMii'li l 'iBt<g!Eyg^t'"'^"'^''' 



LMSC-A604141 



Solve for 1^ 



Roo 



= (K,) (R,i) 



= (1.02100) (2.55) 

= 2. 60 lb oxidizer/lb fuel 



Solve for F„ : 



00 



CD 



= (291.5) (56.81) + 200 
= 16, 760 lb 



The predicted values for flight are determined to be; 



N = 24, 548 rpm 

W = 56.81 

00 

R = 2. 60 lb oxidizer/lb fuel 

CO 

F = 16, 760 lb 

00 

l" = 29 1. 5 (assumed nominal value) 

spoo 

These values reflect typical and not actual predicted performance for 
any vehicle. 



6-22 




LOCKHEED MISSILES & SPACE COMPANY 



«;-*^'gG3»?W5?^^(ff5^'l?W^!*'.^^ 



LMSC-A604141 



6.9 ROCKET ENGINE PERFORMANCE ANALYSIS (POST-FLIGHT) 

In order to verify preflight data predictions, actual flight data are averaged 
over total steady-state burn tinne and analyzed to determine actual values 
for the following parameters: 



• W„ 



R. 



00 



• N. 



03 



From these actual values, the previously assumed I may be calculated. 

'■ ' sp 



6. 9. 1 Methods of Analysis 



The following data are obtained from flight: 



so 



3 
I 
I 



sf 



so 



sf 



"o 

*f 

Pvf 

p 

vo 
OVIP 



Oxidizer suction pressure, average value (psia). 
(Measuremient B-2 Oxidizer Pump Inlet Pressure) 
Fuel suction pressure, average value (psia). 
(Measurement B- 1 Fuel Pump Inlet Pressure) 
Oxidizer suction temperature, average value ( F). 
(Measurement B-32 Oxidizer Pump Inlet Temperature) 
Fuel suction temperature, average value ( F). 
(Measurement B-31 Fuel Pump Inlet Temperature) 
Specific gravity oxidizer (Fig. 3-1 ). 

Specific gravity fuel (Fig. 3-1 ). 

Fuel vapor pressure (psi) (Fig. 2-22) 

Oxidizer vapor pressure (psi) (Fig. 2-22) 

Average oxidizer venturi inlet pressure (psia) 
(Measurement B-11) 



6-23 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



FVIP = Average fuel venturi inlet pressure (psia) 

(Measurement B- 12) 
N = Average turbine speed (rpm) 

(Measurement B-35) 

P = Average thrust chamiber pressure (psia) 

c 

(Measurement B-91) 
= Velocity head of oxidizer (constant 2, 1 psia) 

= Velocity head of fuel (constant 1. 5 psia) 

= Total pressure of oxidizer (psia) 

= Total pressure of fuel (psia). 



velo 



velf 



to 



tf 



The following equations are utilized in solving for the previously assumed 
parameters: 

a. Equation 1: 

P • = P + P 1 
to so velo 

^* Equation 2: 

P . = P ^ + P ,. 
tf sf velf 



c. Equation 3: 



to ^o ^ So / I N , 



where: 



H = head at oxidizer pump inlet (feet) 

so 



6-24 



LOCKHEED MISSILES & SPACE COMPANY 



psjw^Sffwisa^i^ii^v^rrT^ 



LMSC-A604141 



d» Equation 4: 

SI 



(^tf - ^f' 



2. 3l\ /24, 000 



'f 



N 



where: 



H , = head at fuel pump inlet (feet). 
sf 



The oxidizer and fuel pump inlet characteristics, C^ and C^ , are determined 
from Figs. 6-12 and 6-13. 



e. Equation 5: 



W^ = (K.) (C^) ib) (N) 
tco loo 



where: 



W = oxidizer flow rate to thrust chamber 

tco 

K, = constant (to be defined) 



f. Equation 6: 



^tcf " ^^6^ ^"^f^ ^*f^ ^^^ 



where: 



W ^ = fuel flow rate to thrust chamber 
tcf 

K/ = constant (to be defined) 



6-25 



LOCKHEED MISSILES & SPACE COMPANY 



^ir''*^aP^W||P!WnB!W!fW?»^ 



«5WJ5^H^lf???;,V'«?Sll^*'J''-',"iM> "^ *i?Jfl'? 1^1 



LMSC-A604141 































1 








Z 

o 

Ou 










, — DESIGN 


J 










\ 








\ 










^ 





UJ 

oi: 

D 
^/> 

Q. 

o 

a. 

< 
> 

> 
a O 

< 

H 
UJ 
lU 

u. 



o 

< 

UJ 

I 
I- 

UJ 



0. 



S 



o 



M01.-J NOISaa i31NI dwnd/MOTd i3lNI dWnd'^'O^ 



6-26 



4-> 

cn 

•H 

(D 

•4-> 

O 

u 

o 



> 
ci 
O 

o 

1—4 



P. 

In 
(U 
ts! 

• H 

T^ 

•H 

X 
O 

(^a 

■— ( 
I 



■ I-l 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 











































, 










1 

r- DESIGN 












\ 










V 



















UJ 


to 


a: 


• M 


o 3 


M 


CN l/> 


<u 


UJ 

a. 




Q. 


rS 


O 




U. 


A 


< 
> 


O 


o l" 


a 


o > 


o 


'- O 


.f-< 


CD 


+j 


< 


nJ 




4-> 


1- 


• r-< 


Ul 


> 


UJ 

u. 

\ 


O 


o 5; 


^ 


00 3-^ 


o 




1—4 


Q 

< 


h 


UJ 




I 


S^ 


1- 


B 


UJ 


ji 


z 


cu 


°% 


r-l 


<1) 


3 


?J 


D. 


h 




ro 




1—1 




-X> 


o 




v 


, 




tUD 




• r-« 




h 



MOTJ Noisaa laiNi di^nd/Monj lanNi diNnd~'3 



6-27 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



g. Equation 7: 



^ggo = (K^) V(OVIP) (6^ - 0.003) 



where: 



^ggo ~ ^°^ ^^^^ °^ oxidizer to gas generator 



K 



2 - oxidizer venturi constant, obtained from BAG Engine Lo 



gme Log Book 



h. Equation 8: 



W 



ggf 



= (K^) V^ 



(FVIP) (J _ 0.004) 



where: 



^ggf " ^^°^ ^^^'^ °^ ^"^^-^ ^° g^^ generator 

K^ = fuel venturi constant, obtained from BAG Engine Log Book 



i . Equation 9: 



GW = w ^ + W 

^ ggf ggo 



where: 



GW 



rp - total gas generator flow rate (average) 



6-28 



LOCKHEED MISSILES & SPACE COMPA 



NY 



W|IJwSfWI«» II ■I.I 11(^11 jjiup 



i^W » t^-.WJ-'-!l;'.4,viii i i.,iii»im 



J?i.*y-P'"Mi>j|,j»;,«iiW¥i*iy,«*l»i..,.', J! ,7'. 



Er^?OTT*-»t^rt3?Pftr<?^«fl'T?J5PJfWfff^5Bff>«5SIW^^ 



^■=r-. v^ * v- .*'"^> t ■^'^'^' ■ ■ 



LMSC-A604141 



j • Equation 10: 



t tcf tco 



where: 



TCW = total thrust chamber flow rate (average) 



k. Equation 11: 



Wco = GW^ + TCW^ 



1 . Equation 12: 



W 



R 



ot 



00 



w 



ft 



where: 



W , = W, + W 
ot tco ggo 



ft tcf ggp 



m. Equation 13: 



F. = (P^) (C^) ^\) + ^exh 



where: 



exh 



thrust coefficient equals 1.829. This value is a constant and 

was obtained from a special test program. 

throat area (in. ^). This value taken from BAG Engine Log Book 

turbine exhaust thrust. This value is constant (200 lb). 

6-29 



" LOCKHEED MISSILES dc SPACE COMPANY 




LMSC-A604141 



n. Equation 14: 



00 



sp 



W, 



CO 



Definitions of the previously used constants (K and K.) are necessary: 



K, — Defined: 



W 



K, = 



ntco 



{« ) (N ) (C ) 

on n on 



where: 



W 



ntco 



'on 



N 



n 



on 



mass flow rate of thrust chamber oxidizer at nominal 
rated conditions. This value is obtained from the BAG 
Engine Log Book, 
specific gravity of oxidizer at 60°F ( 1. 570) 

nominal turbine speed (24, 000 rpm) 

nominal pump coefficient, oxidizer (1.000) 



K^ -Defined: 



W 



K, = 



ntcf 



(6. ) (N ) (C, ) 

fn n fn 



where 
W 



ntcf 



mass flow rate of thrust chamber fuel at nominal rated 
conditions. This value is obtained from the BAG Engine 
Log book. 



6-30 



LOCKHEED MISSILES & SPACE COMPANY 

!^^''!'^^!^l^l)>l^!l^^:,>.J»j;>lltfil l. j^i.^ 



:^t ^i > jt»^i,t i rt j ^i ' ;.»i'. ?^f»w'^«i?'."'^^^'^ 



LMSC-A604141 



8, = specific gravity of fuel at 60 F (0. 795) 

C- = nominal pump coefficient, fuel (1.000) 



6. 9. 2 Example of Flight Analysis 



Flight Data (Average Value) 



Measurement Giving Values 



p 

so 


— 


14. 30 psia 


^sf 


= 


15. 40 psia 


T 
so 


= 


51.5°F 


sf 


= 


53.0°F 


«o 


= 


1.577 


«f 


= 


0.799 


p 

vo 


= 


1.60 


^vf 


= 


1. 50 


OVIP 


1, 040 psia 


FVIP 


1,005 psia 


Noo 


= 


24, 393 rpm 


P 
c 


= 


504 psia 


P , 
velo 


= 


2. 1 psia 


^,rc.-\f 


= 


1. 5 psia 



B-2 

B-1 

B-32 

B-31 

Fig. 3-1 

Fig. 3-1 

Fig. 2-22 

Fig. 2-22 

B-11 
B-12 
B-35 
B-91 

Constant 
Constant 



6-31 



LOCKHEED MISSILES & SPACE COMPANY 



felfWIOf!!f!W?!P<^9Sy^r^?59^»^ 



rTpff^r' i;yj'iwv fJ'T''r^»'ff^'j* * '' ^ ' *v r^«g'^jS''3;wi 



igifq'-^tgy tl^ lja ^ia^w, ' • 't .^ r^^r*?!^^^?'^?'^-'^ 



LMSC-A604141 



6.9.2.1 Solve for N^, W„ , R^, and F 



CO 



Nqo - Solved: 



N 



(X) 



- average measured data obtained from flight records 



(Measurennent B-35) 
N^ = 24, 393 rpm 



W33 - Solved: 



H 



so 



to vo 



/ 2.31 \ / 24, OOP 



P, = P + P , 

to so velo 



P. = 14.30 + 2. 10 
to 



P^ = 16.40 
to 



H 



so 



= (16.40 - 1.60) (^4^) (^ 



24. 000 



24, 393^ 



H = 20.99 ft 

so 



From Fig. 6-12, C = 0.995 



W. 



too = ^^1^ <S^ (*o) (NJ 

W, 



K, = 



tcon 



1 ~ (6_) (N ) (C ) 
on n on 



6-32 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



where: 



W = 38. 76 (from BAG Engine Log Book data) 

tcon 

38.76 
^1 = (1.570) (Z4,00U) U) 



K = 0.001029 



W 



tco 



W 



tco 



(0.001029) (0.995) (1.577) (24,393) 
39.34 lb/sec 



H 



sf 



^^tf - ^vf 



it) (^; 



H 



sf 



H 



sf 



Ptf = ^sf ' ^elf 



P = 15.40 + 1. 50 



^tf = ''-'' 



(16.90 - 1 
43.09 ft 



'^^> (0.799J \^24,393J 



From Fig. 6-13, C^ = 0.995 



W 



tcf 



= (K^) (C^) (5f) (N) 



^6 = 



tcfn 



(8.J (N ) (C, ) 



fn n 



fn' 



6-33 



w»t !l ^» / ' t "}m '''''''''''^''' "y' ''° S**WW''a ' ■ »* 



LOCKHEED MISSILES & SPACE COMPANY 



"4 "tw^"'*^***^-*' 



■r^-^^rig a - k i r yi»)" »»'«qB>C:-^m — sTrw'cpw^anp 



-<»»wiW!bWJ"""' •IVJ*""'"^*' 



LMSC-A604141 



where: 



^tcfn "^ ^^' "^^ (^rom BAG Engine Log Book data) 



K, 



13. 75 



(0.795) (24,000) (1) 



W 



tcf 



W 



tcf 



K, = 0.000721 



(0.000721) (0.995) (0.799) (24,393) 
13.95 lb/sec 



/^ 



\go = ^2 V(OVIP) (6^-0.003) 



K^ = 0.004650 fronri BAG Engine Log Book 
(oxidizer venturi constant) 



W 



ggo 



(0.004650) \/(l,040) (1.577 - 0.003) 



W =0. 188 lb/sec 

ggo 



W^gj = K^ V(FVIP) (6^-0.00 



4) 



K_, = 0. 045060 from BAG Log Book 
(fuel venturi constant) 



W 



ggf 



= (0.045060) vMiToosMoTTgg'T'oToo 



4) 



w 



ggf 



1. 273 lb/sec 



6-34 



-■''^.waK I " 'uumifgifiifaff^ 



LOCKHEED MISSILES & SPACE COMPANY 



W!ag^^TOgJS »!»!t,ipj ^ j)i^imM . T ;";? y iJ ^ ^ ,. 



LMSC-A604141 



Solving for total flow (W^,): 



Wco = W^ + W 

tc gg 



W^^ = 39.34 + 13.95 



W^^ = 53. 29 



^ee = 1.273 + 0. 188 

W = 1.461 

gg 



W, 



OD 



= 53.29 + 1.461 



W„ 



54.75 Ibm/sec 



Rqo (System) - Solved: 



R 



CO 



W 



ot 



W 



ft 



W = W. + w 

ot tco ggo 

^ot "" ^^'^^ + °- ISS 



W^^ = 39.53 



W 



ft 



W 



ft 



tcf ggf 

= 13.95 + 1.273 



W 



ft 



= 15.22 



6-35 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



R, 



39. 53 
15. 22 



R„ = 2. 60 lb oxidizer/lb fuel 



Fqd - Solved: 



F^ = (P^) (C^) (A^) + 200 



00 



00 



where: 

2 
A = 17. 17 in. from BAG Engine Log Book 

(504) (1.829) (17. 17) + 200 

15,828 + 200 

16,028 Ibf 



6. 9. 2. 2 Solve for new I 



sp«) 



CO 



spec 



Wa 



r— _ 16, 028 



spt 



54.75 



Ibf 



I = 292.75 

Sp03 ,, / 

■^ Ibm/sec 



These values and the analysis are presented as an example and reflect 
typical rather than specific flight performance. 



6-36 




LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



6. 10 ESTIMATED ENGINE ALTITUDE PERFORMANCE 

Table 6-2 is presented to show typical altitude performance and overall 
engine characteristics. 

Table 6-2 

ESTIMATED ENGINE ALTITUDE PERFORMANCE 
AND OVERALL CHARACTERISTICS 



Parameter 


Value 


Characteristic Length 




Thrust Chamber 


83 in. 


Gas Generator 


98 in. 


Characteristic Velocity- 




Thrust Chamber 


5, 220 ft/sec 


Gas Generator 


3, 375 ft/sec 


Flow Rate 




Oxidizer 




Thrust Chamber 


39. 12 lb/sec 


Gas Generator 


0.205 lb/sec 


Fuel 




Thrust Chamber 


13.90 lb/sec 


Gas Generator 


1.370 lb/sec 


Hydraulic Motor Pump 


6. 75 gpm 


Mixture Ratio 




Engine 


2.57 


Thrust Chamber 


2.81 


Gas Generator 


0. 150 



6-37 



LOCKHEED MISSILES & SPACE COMPANY 



r^T'pr^z^'^vpm,'^ 



T }-^3x*'?!!K-''e^ii^9s,'''^fy^T'f^ 



JMiMU^iUSttSCk 



LMSC-A604141 



i 



Table 6-2 (Cont) 



Parameter 


Value 


Pressure (Static) 




Thrust Chamber 


506 psia 


Thrust Chamber Nozzle Extension Exit 


1. 87 psia 


Gas Generator Chamber 


485 psia 


* Turbine Manifold 


475 psia 


Turbine Manifold Nozzle Exit 


12. 6 psia 


Turbine Exhaust Exit 


7. 8 psia 


Specific Impulse 




Engine (minimum) 


291.5 ^^^ 

Ibm/sec 


Thrust Chamber (minimum) 


296. 5 ^^^ 

Ibm/sec 


Thrust 




Engine 




Normal Steady State 


16,000 


Thrust Chamber 


15,800 


Thrust Coefficient 




Engine (Thrust Chamber 45:1) 


1.829 


Turbine Pump Speeds 




Turbine 




Normal Steady State 


24, 800 rpm 


Oxidizer Impeller 


14, 410 rpm 


Fuel Impeller 


25, 389 rpm 


Velocity 




Thrust Chamber Nozzle Extension Exit 


8,727 ft/sec 


Gas Generator Throst 


575 ft/sec 



6-38 



LOCKHEED MISSILES & SPACE COMPANY 



I 



.,lV»M f j;j|; ig>^gyi| «^i^j jl gjUj,».j ii M4jjtji^^ 




^J*S?^i" 



LMSC-A604141 



Table 6-2 (Cont) 



Parameter 



Area 

Thrust Chamber Throat 

Thrust Chamber Nozzle Exit 

Thrust Chamber Nozzle Extension Exit 

Turbine Exhaust Duct Exit 

Turbine Manifold Nozzle Throat 

Turbine Manifold Nozzle Exit 
Leakage 

Pump Primary Seal 

Fuel (nnaxinnum) 

Oxidizer (maxinnum) 

Pilot Operated Solenoid Valve Overboard 
(maximum) 

Pressures 
Fuel 

Pump Suction (total pressure) 
Pump Discharge 
Gas Generator 

Hydraulic Motor Pump Without Orifice 
Thrust Chamber 
Gas Generator Injector Manifold 
Drop Across Gas Generator Injector 
Thrust Chamber Injector Manifold 
Drop Across Thrust Chamber Injector 



Value 



17. 12 in. 
227. 4 in. 
770.7 in. 



12.6 



m. 



0.058 in. 
0. 283 in. 



500 cc 
500 cc 



85 



24 



cc 



psig 



1,045 


psia 


1,045 


psia 


750 


psia 


615 


psia 


130 


psi 


600 


psia 


94 


psia 



6-39 



LOCKHEED MISSILES & SPACE COMPANY 



'A'JWWH9S!?¥^''fc''! '■'■ 'V.?i"»''^r"Ratfa»j 



«*B^PB9rB«?P3«»nr' 



?l...lJJi.lWI-fLil.. 




LMSC-A604141 



Table 6-2 (Cont) 



Parameter 


Value 


Oxidizer 






Pamp Suction (total pressure) 


24 psia 




Pump Discharge 






Gas Generator 


1, 105 psia 




Thrust Chamber 


950 psia 




Gas Generator Injector Manifold 


570 psia 




Drop Across Gas Generator Injector 


85 psia 




Thrust Chamber Injector Manifold 


600 psia 




Drop Across Thrust Chamber Injector 


95 psia 




Speed (rpm) 






Turbine Critical Speed 


30, 750 rpm 




Temperature 






Hot Gas 






Gas Generator 


1,400°F 




Thrust Chamber 


4, 450°F 




Skin 






Gas Generator (Chamber) 


200° to 600°F 




Thrust Chamber (Convergent Nozzle) 
(maximium) 


590°F 




Thrust Chamber Nozzle Extension 
(maximum) 


2, 300°F 




Turbine Manifold 


1,000° to 1, 200 


°F 


Fuel 






Pump Inlet 


60°F 




Gas Generator Injector 


70°F 




Thrust Chamber Injector 


70°F 




Oxidizer 






Pump Inlet 


60°F 




Gas Generator Injector 


65°F 




Thrust Chamber Injector 


65°F 






»#«fc'iii^s».. 



Mjjr— ^'^Wa--^-^-* , 



LOCKHEED MISSILES & SPACE COMPANY " 



LMSC-A604141 



Section 7 
PYROTECHNICS 

7.1 INTRODUCTION 

Ordnance devices used in the Gemini-Agena Target Vehicle (ATV) perform 
some of the most critical flight functions. Ballistic units have been chosen 
because of the history of high reliability and rapid actuation time inherent in 
developed ordnance items. 

7.2 DEVICES USED 

The pyrotechnic devices used in the Gemini ATV are listed in Table 7-1. 
7.2.1 Pressure Cartridges 

Two types of pressure cartridges, LMSC-1062363 -3 (M-69) and 1463174-1 
(M-11), are employed in the Gemini Program to actuate pyrotechnic devices. 
Functionally, the two units differ only in pressure pulse output. The method 
of construction is slightly different in each case. 

Cartridge LMSC-1062363-3 (M-69) —The M-69 is capable of producing 
6000 psi minimum in a 5-cc container and is employed in the first operation 
of the helium control valve. 

The cartridge is constructed with a modified AND 10056-5 threaded output 
end, 0. 559-in. long. It is approximately 1.09-in. long, 0. 5-in. in diameter, 
and has a 0. 625 in. -diameter hexagon in the center for wrench flats. The 
connector end is manufactured to nnate with the Bendix Pygmy PT06E-8-4S 
bayonet type connector. Internally, the M-69 has four connecting pins 



7-1 



LOCKHEED MISSILES & SPACE COMPANY 



Table 7-1 
GEMINI PYROTECHNICS 



LMSC-A604141 



LMSC 
Part No. 



1062363-3 
1463174-1 



1395698-1 

1396224-501 

1347107-503 

1312289-503 

1062410-5 



1354376-501 
1461809 



Nomenclature Quantity 

Standard Agena D 
Type M-69 Squib 2 



Type M-U Squib 



Separation Charge 
Separation Detonator 
Self Destruct Charge 
Self Destruct Initiator 
Booster Retrorocket 

Progrann Pe culiar 

Separation Bolt Assembly 
Detonating Cartridge 



12 



1 
1 
1 
1 
2 



4 
8 



Where Used 



POHCV* 

POHCV-, Horizon 
Sensor Panels 
Pin Pullers and 
Pin Pushers 

Booster Adapter 
Booster Adapter 
Booster Adapter 
Booster Adapter 
Booster Adapter 



For Separation 
Bolt Assembly 



*POHCV — Pyrotechnically operated helium- control valve 

(to match the Bendix connector) contained within four individual glass seals 
fused to individual holes through the cartridge body. A 0.003-in. -diameter 
nichrome bridgewire is soldered across each set of two pins, A to D and 
B to C. A bridgewire priming mix, composed of lead styphnate and a 
desensitizer additive, encapsulates the wire ends and is subsequently 



7-2 



LOCKHEED MISSILES & SPACE COMPANY 



w-^BSpMrpr"-"--- ''--^■^-f'«>'si » ' j BW Y ^' N !iffy"- r 'j' '.B'S ' ji .i-^ ^ y" 



lpM5J»W|j»^¥WV5f«»-i!5!««((rt!i'- 



LMSC-A604141 



surrovmded by a circular phenolic sleeve. Approximately 227 mg (3-1/2 
grains) of Hercules Bullseye powder are loaded on top of the priming mix to 
form the main charge. The output end of the cartridge is sealed with a 
0. 006-in. -thick lead closure disc which is cemented in place with NP 428 
adhesive. The cartridge end is crimped over to retain the seal assenably and 
then coated w^ith Poly-Ep-Platon epoxy compound to complete the closure 
process. Table 7-2 presents the electrical characteristics of the cartridge. 

Table 7-2 
CARTRIDGE ELECTRICAL CHARACTERISTICS 



Maximum Test Current 

Resistances: 

Bridge Circuit 
Any Pin-to-Case 
Bridge -to- Bridge 

Maximum Safe Current 

All-Fire Current 



10 mia per bridge 

0.45 to 0. 85 ohm at 70. 0° ±20°F 

1 meg minimum 

1 meg ininimunm 

0. 5 amp per bridge for 5 min 

@ 26.0 ±2.0 vdc 

5.0 cinnp per bridge for 15 ms 

@ 26.0 ±2.0 vdc 



Cartridge LMSC-1463174 (M-11) — The M-11 cartridge is capable of producing 
3000 psi minimum in a 2-cc closed volume of 70° ±20°F. This M-11 cartridge 
is used in the second operation of the pyrotechnically operated helium- 
control valve (POHCV). The M-11 cartridge is constructed with a modified 
AND 10056-5 threaded output end, 0.300-in. long. The tapered portion of 
the standard AND has been removed to allow the output end of the cartridge 
to be efficiently sealed. This cartridge is approximately 0.840-in. long and 
0. 5-in. in diameter, and has a 0. 625-in. -dianmeter hexagon in the center for 
wrench flats. The connector end is manufactured to mate with the Bendix 
Pygmy PC06E-8-4S screw-on type connector. Bridgewire connecting pins are 
sealed w^ith separate glass seals fused to individual holes in the cartridge body. 



7-3 



LOCKHEED MISSILES & SPACE COMPANY 



^ftt,^»;j^ » iy^ j^jy^^^A T w>v . ^ .' J ' ■. ' , [.^ ' JU{ m^ »9 ^F^^rr'■ryv?:^''!fsrT.'; 



T? gt% y:''tr';^^-^Tyc.?'^^ ^ '* :' <B s y . '; yjt 'wCT"*rn^'' ^-t.': 



LMSC-A604141 



A 0.003 -in. -diameter nichrome bridgewire is soldered across each set of 
two pins, A to D and B to C. A bridgewire primary mix, composed of 
zirconium and barium chromate, surrounds each of the two wires, and 65 m^^^ 
of Hercules Bullseye powder are loaded around and on top of the primary mix 
to form the main charge. The output end of the cartridge is covered by a 
0. 006-in, -thick lead foil closure sealed in place with a low -melting -point 
solder. Table 7-2 gives the electrical characteristics of the cartridge. 

7.2.Z Booster Adapter Separation 

Separation Charge — The separation charge provides the energy to fracture 
the booster -Agena joint. This separation charge is a mild detonating fuse 
(MDF), which is a 10 grain per fbot RDX explosive core, contained by a 
flexible lead-antimony sheath and covered by a thin polyethylene extruded 
coating. The MDF is used in the separation system because of the following 
highly desirable properties: 

1. High detonation rate, approximately 20, 000 fps which guarantees 
almost instantaneous separation 

2. Controlled predictable explosive power which prevents inadvertent 
damage to surrounding components or structure 

3. Relative insensitivity to shock friction, heat, and stray electrical 
energy which reduces the hazards in handling, storing, and 
installation. 

The RDX core is identified by a maximum of 1 -percent pink dye molded into 
the material, which provides for easy detection and disposition in the event of 
any inadvertent spillage. Sheathing alloy is composed of 7 -percent antimony 
in a lead base to increase the strength, and a polyethylene cover to provide 
waterproofing in case of sheathing porosity. 



7-4 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Fracture Joi nt — A frangible ring section is constructed with three closely- 
adjacent parallel grooves milled around the inside circumference of the 
booster adapter. In addition, three longitudinal grooves which cut across the 
circular grooves are placed 120-deg apart around the inside of the ring 
section. Detonation of the fuse fractures the joint, thus producing a gap 
between the booster adapter and Agena, Six small strips of the adapter 
material are blown out and away from the vehicle. 

Detonating Cord Ring —An extruded -aluminum detonating -fuse ring is fastened 
over the three parallel grooves. This ring contains the fuse, provides backup 
for the explosive reaction, and focuses the detonation pressure on the grooved 
section of the frangible joint. 

Separation Detonator — The separation detonator assembly consists of two 
LMSC-1062876 electric detonators positioned in a detonator holder and re- 
strained by Locktite sealant compound. These detonators receive a firing 
stimulus which heats up a lead-styphnate -coated bridgewire. The lead 
styphnate reacts to the heat produced by the bridgewire and detonates; this 
reaction is transferred to the adjacent lead azide which also detonates and 
initiates the main charge of RDX. Firing of the two detonators initiates the 
firing of the ends of the detonating fuse protruding through the collets. 
Table 7-3 presents the electrical characteristics of the detonators. 

Table 7-3 
SEPARATION DETOxNATOR ELECTRICAL CHARACTERISTICS 



Maximum Test Current 
Resistances: 

Pins A to D 

Pins C to B 

Pins A to C 

Pin A to Case 
Maximum No-Fire Current 
Minimum Sure-Fire Current 



10 ma at 50 vdc 

0.90 to 1. 10 ohms 
0.90 to 1. 10 ohms 
1 nneg minimum 
1 nrieg minimum 

0. 50 amp for 5 min 

1. 5 annp for 6. 5 min 



7-5 



LOCKHEED MISSILES & SPACE COMPANY 



B<iy. MiBffiWi! W i > l|; .y i .tJ..yi| ( iyj i K. fj y) j pM^j^ i ^ 



F^'^TfW^^^wpWS' 



Lv sm.^y.iw g ; w ?3 ? |f ^ y ; ^ {i ^ M r y -'^ 



LMSC-A604141 



7. 2. 3 Vehicle Destruct Systems 

The vehicle incorporates two separate destruct systems: self-destruct and 
command destruct. The self-destruct system provides a means of destroying 
the vehicle during the boost phase of flight, in the event that an uncontrollable 
condition occurs wherein continued flight would constitude a hazard to property 
and personnel. The command destruct system is used only for control 
purposes. In this application, the pyrotechnic devices are removed from the 
system. The command receiver and associated equipment a re used to transfer a 
shutdown signal to the engine. 

7.2.3. 1 Self-Destruct System . This system consists principally of a shaped 
explosive destruct charge, a destruct charge initiator, four auxiliary destruct 
batteries, two premature separation switches (also known as destruct lockout 
switches), and related destruct system circuitry. All system components are 
installed in the satellite adapter section. The destruct charge, initiator, and 
batteries are adjacent to each other on the outside of the adapter, just aft of 
the vehicle separation plane. The two premature separation (destruct lockout) 
switches are inside the adapter. 

The self-destruct system is capable of being actuated by a command signal 
from the Range Safety Officer or auto naa tic ally upon premature separation of 
the Agena-D vehicle and booster. This system is electrically connected with 
the first-stage booster destruct system to provide the capability for simul- 
taneous destruction of both vehicles by the ground command signal transmitted 
to the booster destruct system. Destruction may be effected any time between 
liftoff and booster thrust cutoff. Safety features are incorporated into the 
system to prevent accidental detonation of the destruct charge (1) during pre- 
launch checkout on the launch pad and (2) during flight until the normal 
separation of the Gemini-ATV and booster. 



7-6 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



Shaped Charge — the flight termination system includes an explosive shaped 
charge (Figs. 7-1 and 7-2) and destruct initiator (Fig. 7-3). This assembly 
(Fig. 7-4) is mounted in the booster adapter and so aligned that the center - 
line of the charge and initiator assembly pass through the standpipe in the 
propellant tank and the common diaphragm separating the fuel fromi the 
oxidizer. 



Destruction of the Agena vehicle results when the tank stand pipe is severed 
and the diaphragm is ruptured, thus allowing the fuel and the oxidizer to mix. 
This is accomplished by firing the shaped or coned charge. Since the pro- 
pellants are hypergolic, vehicle destruction is instantenous upon mixture. 
By using this method, in which a metal cone is backed by a high explosive and 
is concentrated or focused into a metallic jet on firing, the necessary pene- 
trating power required to pierce the tank, penetrate the oxidizer, and rupture 
the diaphragm is achieved. 

The shaped charge assembly is 3.600-in. in diameter and 5.33-in. in length 
and weighs 8. 5 lb. Of this total weight, approximately 1. 3 lb is composition 
A-3 high explosive. The dual booster pellets are cased RDX, 0. 590-in. 
diameter. RDX is also the explosive m.aterial used in the relay pellet, which 
is 1.480-in. in diameter and 0.75-in. thick. All the explosive components — 
charge, booster, and relay —are pressed items which are bonded into the case 
with an epoxy connpound. The copper cone, w^hich has a 60-deg angle with a 
0. 580-in. diameter flat at the apex, is also bonded to the charge. The edge 
of the case is crinnped over to hold the whole assembly securely in place. 

Initiator — The explosive initiator assembly (Fig. 7-3) contains three fuse- 
train elements. The first element is separated from the third by a steel 
rotor; this rotor contains one hole which either aligns with (arms) or dis- 
aligns with (nnakes safe) the explosives. The rotor hole contains the double - 
end second elenaent. When this elenaent is rotated to the arm position, the 
explosive train frona the first to the third element is completed. 



7-7 



■■ LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




W 

a 
o 

U 



0) 
tUD 

yi 
ri 

M 

U 

a 
CO 

a 
u 

•H 

a< 



I 



7-8 



LOCKHEED MISSILES & SPACE COMPANY 



l^»^i i )t:». » ; | t>«»y i ! |i |fi 1 ii i )i!ii^j^iij^j<.>w»jV' ' 'tMWt^^^^^^^ 



■7 l ". m;, 'm. Kmi'f.,:r:ktK"H ly »Vl^ ' ^■ ^■'nryrl ^^^.l,.^J^y,J^Hl ^! ,.^■^..^^;-J » ';^^^^> ! >!a^ 



LMSC-A604141 






/ 



\ 



Fig. 7-Z Shaped Charge 



7-9 



LOCKHEED MISSILES 8: SPACE COMPANY 



ffly | Kj !!!rrP "^**Ji i «* }i i> ff^ '-gi S^*^''WW^'«J ' " "' ^ ' '^)^ ' ^ ' *^^ ■*-m'-%r3ai(f^,^Wti^r5Sfw«5!r,p"rSf*pv SF* »<»««;« "»5*^'W^' ■^^'J*^ ■=r^'««->^0*<5»P«^--«>T»-'ps^^'?*r?w^?^||^ft^ 



LMSC-A604141 



B 



\ 



I ^ ; ;rro y 

i^ if ' ' "-* S p 

1 1 fl. tu f . 5 -1 ! 

II Q I . O -J i ', 

I UJ O h Z ^ ■ • 










«^j.> fry 



tJ jfefcLJ^' 






u 
o 






u 

4-> 

to 

Q 

oo 
I 

h- 

bi) 



e 
I 
I 



7-10 



I 



LOCKHEED MISSILES & SPACE COMPANY 



I 



f^ i uwBfcwi,i'>BMWj i j :!;' V B » y.. ?u i ?iy^i »ji i | i^^ t.-.y 



LMSC-A604141 









"' t ii\ 



\ 



•x.. 



^ 0: 




«-M>*j»^.,^j^_^ 



ut^ 



■^ 



,J — ' 



J. ;s»s^ '-.aP" ■' ~ ■ 






^ f*;^ /r^. 












Jjftj I— < 



ft 



f»»WB5PJW 



■-1 t 

»«.■■ i 






1—! 

a 

M 
W 

< 
U 

o 



ni 

^1 

O 

(15 
CO 

■ H 

P4 



7-11 



LOCKHEED MISSILES & SPACE COMPANY 



^jWljWilW 'W i ^' ^?^^- ' ^ ' ^^'' ' ^^?^ ^ 



'r-t)p5^jB«:*»WK»«ffi«W»tr^;'pnwrP'g^^ 



•'■•F*-®'' T!?T^»'"-'""^"")iffl!l?W"'''H**7 



ggEPft ^y 7 i:j g i.^ ' »J ' "*t¥ .. 4 ^'- " ^^ ' 'j^ gj jyi^^^fflg ' *''" ' ^ ^'^^ ''V^ 



L,MSC-A604141 



A Ledex switch, mechanically linked to the end of the rotor, converts an 
electrical impulse into a rotary motion. A mechanical latching mechanism 
(ball detent) assures that the rotor positively positions itself at 90-deg 
(large initiator) or 180-deg (small initiator) intervals. Repeated activation 
of the Ledex switch successively rotates the rotor 90 deg (safe-arna-saf e-safe) 
or 180 deg (safe-arm-safe) and rotation is always in the same direction. 

To lock the rotor in the safe position, a safety pin is inserted through the 
explosive initiator assennbly and a hole in the rotor. This safety pin is 
reinoved after the destructor has been electrically connected and all vehicle 
checks have completed. 

A window in the end of the explosive initiator assembly permits a visual 
check of the rotor position. 

Arming Circuit — The arming circuit is designed for remotely arming or 
disarming the satellite flight termination systenn. Remote control is from, 
the blockhouse and is accomplished by advancing the rotor-Ledex assembly 
of the explosive initiator to the arm safe positions. Arm and disarm func- 
tions are switch-operated by wafer switches connected to the rotor-Ledex 
assembly. The wafer switches function like a two-position latching relay 
when the rotor assembly is rotated; only the mechanical rotation of the rotor 
actuates the wafer switches. The arming circuit is also isolated electrically 
from the battery-destruct circuit by this arrangement. In the safe position, 
the bridgewires in the explosive detonator are shortened, and the auxiliary 
battery leads which supply the input signal current are electrically isolated. 
In the arm position, the shorts across the bridges are opened and the bridges 
are electrically connected to the input signal leads. 

Electrical Characteristic s — The explosive detonator is provided with two 
bridgewires, each capable of initiating an explosion. Characteristics of the 
detonator are presented in Table 7-4. 



7-12 



LOCKHEED MISSILES & SPACE COMPANY 

t .^J^^ai^||l|l^j ^| |K B >™■*^W'»^^lj* ' "'!^lM i w|l | » '' ^wq^^^g» 



LMSC-A604141 



Table 7-4 

SELF-DESTRUCT SYSTEM DETONATOR 
ELECTRICAL CHARACTERISTICS 

Bridgewire Resistance 0.8 to 1.2 ohms (70°F) 

Maximum No -Fire Current 0.5 amp for 5 min 

Always Fire Current 2 amp at less than 2. 5 



"^•^•^•2 Command Destruct System . In addition to the self-destruct system 
located on the booster adapter, command destruct receivers, which are 
optional equipment available for Agena-D vehicles, have been installed to 
provide a means of permanently shutting down the engine. For Gemini 
vehicles, the explosive charge is removed, leaving only the antennas and 
receivers. The system is actuated upon receipt of a ground command through 
a redundant system of two antennas and two Agena -mounted command destruct 
receivers. The receivers then transmit a pulse through the flight control 
J-box to the aft safe/arm J-box where two pairs of relays are tripped, the 
engine start signal is interrupted, and the engine is shutdown as described 
in Section 4. 

Shutting down the engine with the command destruct system permanently 
disables the Primary Propulsion System, because the relay contacts in the 
aft safe /arm J-box cannot be reset. Therefore, once the vehicle mission 
has been completed, the command destruct system can be actuated and the 
vehicle can not be fired by any spurious electrical signals. 

7.2.4 Booster Retrorocket, LMSC-1062410-5 

Upon actuation of the separation charge, the Gemini -ATV is released from 
the booster adapter. The booster retrorockets provide reverse thrust to 
reduce the booster velocity relative to the Agena velocity thus causing positive 
separation. The 0. 9 KS-500 solid-propellant retrorocket motors, 
LMSC-1062410-5, are designed and qualified by Rocket Power, Inc. Each 
rocket motor develops approximately 500 lb of thrust for a nominal burn 

7-13 



■ LOCKHEED MISSILES 8c SPACE COMPANY 



LMSC-A604141 



time of 0.9 sec. The maximum length of each unit is 15.25 in. and the 
maximixm diameter is 2.9 in. (See Fig. 7-5 for a cutaway view.) The motor 
case is a cylindrical assembly of AISI 4130 alloy-steel, having a minimum 
yield strength of 135, 000 psi. Nozzle and forward closure assemblies are 
also of AISI 4130 steel with a graphite insert in the nozzle throat. For 
thermal control, the case, forward closure, and the outside of the nozzle 
are either coated with white epoxy paint or nickel plate. Four machined and 
drilled lugs, welded to each motor case, are used for attaching the rockets 
to the booster -adapter mounting pads. The rockets are mounted 180 deg 
apart on the Z-Z axis, with the nozzles canted away from the booster adapter; 
the centerline of thrust runs through the empty CG of the booster. Mounting 
provisions allow for alignment of thrust with respect to the booster em.pty CG 
and adjustmient of motor case to X-X axis parallelism. See Table 7-5 for 
performance characteristics and Table 7-6 for weights and dimensions. 

Table 7-5 
PERFORMANCE CHARACTERISTICS OF 0.9 KS-500 RETROROCKETS 



Propellant 

Duration 

Thrust 

Chamber Pressure 

Total Impulse 

Ignition Delay 

Operating Temperature Range 

Storage Temperature Range 

Auto -Ignition Temperature 
of Propellant 

Maxinnum Guaranteed 
Ignition Altitude 



2. 1 lb RPI-PAP-8 polysulfide fuel and 
amnnonium perchlorate oxidizer 

0.925 sec nominal 

490 lb nominal ' 

725 psia nominal 

455 lb-sec ' 

0.013 sec''' 

-40" to +200 "F 

-20° to +120°F 

360°F 
150, 000 ft 



•'At sea level and +70°F. 



8 



7-14 



LOCKHEED MISSILES & SPACE COMPANY 



l ?- 'y gf ? ^Wj w yt«iy y ;yiitMP i ^ ii |j»>iutiw i , p^^^ ?ff ^ ' -n ^<F?r^*9CT*'gwcaKy7r^^ ?? t^ » yW J? -« ? ! »■ !' >.■■ -f^ ^J ^ ^ !- -. ' - •»'-^'**^r -'i} ^ v th\-'T-x~ ' '<i :' i !t .^ ' ! lf' :3^ - :^ - »rv' ^^^ 



LMSC-A604141 



I 





UJ 



a 
z 

UJ 



W 

(Tj 

O 



O 
O 

o 

• H 

t4 



7-15 



LOCKHEED MISSILES & SPACE COMPANY 



; 'i ?<j> y iw^^t«jai^ i ffji)p»j4iB.;i-iiJiW'jiuijigjajM^^ 



■Si>^Ty;r^« TOWI- i | ^ it y ^ 



^'f"??';**^ 



LMSC-A604141 



Table 7-6 
NOMINAL WEIGHTS AND DIMENSIONS OF RETROROCKETS 



Rocket Motor Dimensions 
Overall Length 
Overall Diameter 
Weight Before Firing 
Weight After Firing 

Nozzle 

Throat Area 
Expansion Ratio 

Rocket Motor Weights 

Motor Case 

Propellant Grain and Liner 

Forward Closure 

Nozzle and Insert 

Igniter 

Miscellaneous (gas seals, diaphragm, 
and inhibitor) 

Total 



15.25 in. 
2.9 in. 
4. 69 lb 
2. 56 lb 



0.455 to 0.458 in. 
7. 6 to 1 



1.20 lb 
2. 16 lb 
0.36 lb 
0.67 lb 
0. 15 lb 
0. 15 lb 



4.69 lb 



Starting from the forward end of the motor, the component parts in order 
are: the igniter, forward closure, gas seal and O-ring, propellant grain 
and liner, motor case, aft gas seal and O-ring, nozzle-closure diaphragm, 
and nozzle assembly with graphite -throat insert. 

The igniter is an electrically actuated hot-gas type. The igniter assembly 
consists of a 1-gm pulverized boron-potassium, nitrate booster charge and 
a 2-gm pressured boron-potassium nitrate sustainer charge, contained in 
a threaded steel shell which is hermetically sealed to the forward closure. 
A double bridgewire ignites the booster charge. The igniter mates with a 
3106-10SL-4S Bendix-type connector. 

7-16 



LOCKHEED MISSILES & SPACE COMPANY 



l i^; a «j ^ ; ■ i-i yn^,< |» . '. M f«.'g''»i'K '■f--^'. 



LMSC-A604141 



The forward closure, including the hermetically sealed threaded-end 
igniter, is screwed to the forward end of the motor, and an O-ring seal, in 
combination with sealant applied to the threads, completes the sealing 
process on the forward end. 

A hot-gas seal, bonded to the forward end of the grain, limits the amount of 
hot gas which is applied to the threads and outer periphery of the forward 
propellant grain face. Due to the design, the sealing ability is improved as 
the pressure increases. 

The solid propellant grain is a polysulfide fuel-ammonium perchlorate 
oxidizer composition with an internal burning, six-pointed star configuration. 
This composition is bonded to the motor case with a polysulfide slurry mix- 
ture which also acts as an inhibitor. The case,, which is threaded on both 
ends to receive the closure and the nozzle, is tested of 3500-psi proof 
pressure. The aft end gas seal is of the same material as the forward seal 
and is bonded to the grain. 



Between the gas seal and graphite -nozzle insert is a nozzle-closure diaphragm 
made of 0.030-in. thick aluminum, which is designed for burst at 150 psi. 
To retain the nozzle insert, the nozzle assembly has a machined well. The 
assembly is threaded to the case and sealed with an O-ring and sealant on 
the threads. 

Electrical Characteristics — The electrical characteristics of the booster 
retrorocket are presented in Table 7-7. 

Table 7-7 
BOOSTER RETROROCKET ELECTRICAL CHARACTERISTICS 

Maximum Test Current 10 ma per bridge 

Resistance: 

Bridge Circuit 0.7 to 1.3 ohms 

Bridge-to-Case 10 meg minimum 

Maximum No-Fire Current 0.2 amp for 30 min 

Minimum Sure-Fire Current 3 ±0. 15 amp for 0. 1 sec 

7-17 



• LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



•4 



7. 2. 5 Pin Pushers 

Pin Pushers are utilized for the purpose of releasing and ejecting the hori- 
zon sensor fairings at vehicle separation. A typical pin pusher can be seen 
in Fig. 7-6. The pin-pusher has a two-piece piston and pin assennbly in an 
aluminum case with pressure cartridges for the two ports. The portion of 
the pin assembly which fastens to the sensor fairing is retained in the 
pusher body by a shear pin. The piston and the other half of the pin assembly 
are mounted in the pin pusher actuating cylinder and positioned by a stop on 
one side of the piston. As the pressure cartridges are fired, the fairing is 
released which pushes the piston, breaks the shear pins, and drives the 
fairing away from the vehicle. Two pin pushers are used on each of two 
fairings, one located at each end. 

7.2.6 Pin Puller 

Figure 7 -7 shows the horizon sensor torque tube actuator pin puller. 
Actuation of this pin puller perm.its the horizon sensor torque tube to change 
the position of the horizon sensors with respect to the vehicle, since the pin 
acts as a positioning stop for the torque tube. Removal of this positioning 
pin, by firing two pressure cartridges in each puller, allows the torque tube 
and sensors to move to the next position by spring action. The pin puller 
has a one-piece piston and pin assennbly inserted into an alunainum body, 
containing tw^o pressure cartridge parts and a naounting flange. The pin is 
drawn into the body by the application of gas pressure to the pin side of the 
piston. 

I 

7.2.7 Separation Bolt 

The separation bolt is used for separation of the protective shroud fronn the 
Agena vehicle. A view of the double-ended bolt with cartridges installed is 



7-18 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




SQUIB PORT 



Fig, 7-6 Separation Pin Pusher —Horizon Sensor Fairing- Vehicle 



7-19 



LOCKHEED MISSILES & SPACE COMPANY 



■^i;' ,.>>k*J-W i* W ! *y ' ^^ . ' * l -*^ '-" wwBBayi ayt/i* ^i|i-i<gl«y-J«> WW 9' ' fmff y ii i |ii^i^iy|L j ffli yt ijT y 'yTrj;?^^:!^?? ^ ^ 



rr^?^gjp^yr?;,? »rr - g^ ^ir! y w'T tn3J*»iMyi ;S.^B^t'J l . | ylJH ' ^l. ^ ^y-^y^.»J^ ' W:i^^l ^ ^ ' W« y B w MJ P ■^^^ ^ ^ ^ 



LMSC-A604141 




1 — I 
■—I 

Oh 
C 

a. 

C 

•r-l 

rt 
:3 
+j 
u 

< 

43 

H 
pi 

O 

H 

O 
in 
C 
(U 
W 

c 
o 

N 

• iH 

^1 

o 
K 

I 
r- 






7-20 



LOCKHEED MISSILES & SPACE COMPANY 



^^^'^^!^M9^'^pv^»t;ifi^f^lf^aif^ 



' -^^'i!iy.' ' ftH'V ! -' i 'I' ' V'- -■ ' : ' ^Mffy w^ ^5|Wt ' .w . yj^^, jy -?'!? ^ jii.- 1^^^ ^ ^ . ijjv w j^ fi^-, ^•. ^■,!i f> .'f^ } s.^Mm^ fffKr^;j^mm^;«f'r^ H-.; ; y f j. ii i jKy;t:iy^>^-y».w^«^w^ ^ yj,,, ,. |,yi j_^ i n^M-.^ 4 ^. jw !^?i i*!y^r' ! *>gr ■iw,4i>,y f^yrgyy " i^-^ -r"?.?r'^» r 



LMSC-A604141 



shown in Fig. 7-8. The bolt is 3. 19-in. long, ll/l6-in. 24 external 
threads, and is made from 4340 heat-treated steel. The bolt carries an 
ultimate load in excess of 16, 000 lb. Upon application of firing pulse, the 
bolt fractures within 10 ms. 



The separation bolt is used primarily as a release mechanism and does not 
provide the energy required for the shroud separation. Fracture of the bolt 
releases a shroud clamp mechanism and allows the shroud to separate. 
Since one fracture is sufficient for clamp mechanism separation, the two 
fracture capability of the double-ended bolts is redundant. Additional redun- 
dancy is provided by the dual bridgewire of the detonating cartridge; eith« 
bridgewire will cause the cartridge to function. 



ler 



ise 



Detonating Cartridge - The detonating cartridge is intended solely for us 
with the separation bolt. Upon actuation, bolt fracture is caused by a com- 
bination shock-pressure effect. The detonating cartridge is constructed 
with a 1.310-in. long, l/2-in. 20-thread output end. Overall length of the 
cartridge is approximately 1. 82-in. The center of the cartridge has a 
0. 625-in. diameter hexagon protrusion for use with flat wrench. The 
connector end is manufactured to mate with the Bendix PC06E-8-4S, The 
base charge is 135 mg of PETN, and the transfer charge is 203 mg of 
lead ozide. The shell is stainless steel. The electrical characteristics of 
the detonating cartridge are presented in Table 7-8. 



Table 7-8 

SEPARATION BOLT DETONATING CARTRIDGE 
ELECTRICAL CHARACTERISTICS 



Maximum Test Current 
Bridge Circuit: 

Pins A-to-D 

Pins B-to-C 
Pin-to-Case Resistance 
Maximum Safe Current 
All-Fire Current 



10 ma per bridge 

0. 5 ±0. 1 ohm 
0. 5 ±0. 1 ohm 

1 meg minimum 

0. 5 amp for 10 min. 

2 amp within 10 ms 



7-21 



LOCKHEED MISSILES ft SPArE r.OMPAMV 



LMSC-A604141 



-I 

-J 

UJ 


o 


X 


UJ 


1/5 


Di 


Q 


3 


UJ 


1- 


Q 


U 


^ 


< 


o 

J 


u. 




o 
CQ 



u 
W 

> 

■ H 

to 

O 

I— I 

X 
W 

00 
I 






7-22 



LOCKHEED MISSILES & SPACE COMPANY 



'■^^T!«^■T-'■.,'?*^y^^?^/TJ^^1 



LMSC-A604141 



Section 8 
AEROSPACE GROUND EQUIPMENT 

8. 1 INTRODUCTION 

Program-peculiar Aerospace Ground Equipment (AGE) is required to support 
Gennini Program peculiar requirements at Santa Cruz Test Base, AMR 
Hangar E, Launch Emplacement 14, and LMSC Sunnyvale. Existing AGE for 
the Model 8096 engine was inadequate for Primary Propulsion System (PPS) 
checkout; therefore, new AGE was designed to meet requirements of the 
Gemini Program. The required AGE is: 

• PPS Checkout Console, along with its acconnpanying tools 
(Fig. 8-1) 

• Start Tank Loading Carts (Figs. 8-2 and 8-3) 

• Auxiliary Flushing Unit (start tank flushing) (Figs. 8-4 and 8-5) 

• Vacuunn Start Tank Bake-out Oven (Fig. 8-6) 

8.2 PRIMARY PROPULSION SYSTEM CHECKOUT CONSOLE 

The PPS checkout console (Fig. 8-1) is aportable pneunnatic- electrical unit 
utilized in perfornaing the following checkouts: 

1. Model 8247 engine functional checks including those for main and 
gas generator (GG) valve actuations, and for start tanks and dual 
check valves 

2. Model 8247 engine leak checks including those for turbopump seals, 
main oxidizer system, main fuel system, thrust chamber, turbine 
exhaust, nnain valves, and GG valves 



i-1 



LOCKHEED MISSILES & SPACE COMPANY 



^ r^ r ' i^Mjv i' ^^tf^ w^ --^-'*i«'*«-7^r?*'^^"?jr*y* - ! ^^ '? w-H •' i ' w i -p ' ^'gwy-gyjwagig^ .T^^v??^'^!5^ws^is«'^.rws*?^ 



LMSC-A60414 1 






» p It * 



ts 



u 

'UJ 



a: 
1 O 






G 



G 



,, U:l O Z 
D. K D_ ' 













i--^ti 






f " ■ 'if 



L... 



//.^.., .i:,-- 


Vj,; 




f -IJ 


X — _^ 


f: 1 


<s^ 


^■. .1 




yd 




111 



Li-iM^J 



r^:3 



1 :^^im.*.,»: !*'^ 










4J> * ' . ' 



,-/^^-.»^»^, B«KC-^.^;. *««^^^ ^ JHEia^y .^. 



~?*»!3P« 



r Mill r'T^jN^Sfatj/il ' I [^, r ajjinfrftl Ji IMMli.^ft t>iJUfcuau.^.»i<^«*i 



O 

c 
o 
U 



u 

o 
X. 
U 



^ 


>^ 


■i 


CO 


1 


c 


i 





i1 


■H 




m 


1 


1 — [ 


i 




a 


cx 



o 

s 

u 
0^ 



00 

•r-l 

hi 



tost 






!3 



^ 8 



8-2 



LOCKHEED MISSILES & SPACE COMPANY 



'^■"5'*"'-"i**^' ■■**?■ 



If ^yr^ggryg-w?' ^5fe.g:^-T?'^T'^^?^"^^'9S???^^S^r^-T'-- 



■^ '^ ^' ^ ' ' y ?r T r . TT^y^^-''^.''''yy'** -'? ' *g * '^-wy * ' ** 



»"flWJ!^igyi^B?f!W?e^^^?;M'^^W!?e^^ 



LMSC-A60414] 



'■■'?^^^m^:>f''K^f-^'fKnr_-;''-^iT> ^j^l«i >' \f l ^K^ ^ 



'■^- T^j i^r ^ ' >■??*■ '''^?<**TTsw'^tjPwv;wjiftj|?.A/;iui^;:(,'«>«Baij|(? ■'^■^.^;^ 



1 L_.. 



Kr::m 



r::.^ 



11 ((>^:) i^o 





t;^^;^ 




'-■^. 


:"/':) 


' ^ ?>; 



L--^ 






v' ■, i 



..J : '^^ 






,^ 



[:..:r:ii 






1 j; J^ 



\S,F^ 



4—-.^' 



>rS33»iWffii^*'' 



' -*T9^a^ 



i 







7 ^M 



„.^?-f sTTPr.- 






E- 



(...) 


















1' fS^^'^^V^f-'" 




\;._wi.;.ii^^S'^ 



Fig. 8-Z Oxidizer Start Tank Loading Cart 



8-3 



LOCKHEED MISSILES & SPACE COMPANY 



T^a- ' w y g ' J,* ff -g^' .lJ^ l W;:w. g' gv ■fy J y j < ^ ^ r ' . * ^ T?y ' « g ^^ » w^ ''?? j ' -, ?*»Hgy?qi^ '*j l w ^ *^ y p ^ 'y■ yyL^3^J ! ^^iCT ^^:f w is ^. tw ' ^vw g ygy^ ^g ?wg'g*^'ry ! ^ !F^ ■ ^y^g^j i! ^^y^ ^ ^«^ 



LMSC-A604141 



"■""•^^^sgSS??^' -r^^^p ''^, 






1! I 









f^na 









ES!:rg 



/■ 



/ 



..Su-." V-,? { 






r 



f' ■ ..'SIS 



.X 



'■C^x E^33 ' / ^^'^ KS-33 









a 



i^/".; 

:---I:i/ 



I i 



Lw'-:I-',-*£ 






K«?3 













«'■ 






■ --.ii 



'/ "V 



I s 



^>^"js.. 



X 



IWf ^fl't-^' ■; 






1 I i 



J" 






\^ 



-- <^-.JW. 



>;:...,..--»-* 




1 
I 



Fig, 8-3 Fuel Start Tank Loading Cart 



8-4 



LOCKHEED MISSILES & SPACE COMPANY 



^ ^ ^ii i w,.fjffg ia ffiy. . wtypjy^,^w |[ ^yjn i|. jjt i. ^ HT.111- j^mm^f T^-- ^-^ ^y.v«t i w y y igp? f.>s^(^^.ui^,: ' if^ ! ^- '? - ! *'y.- ! g^ ' ' ' g g^ 



!*W^'?E5^^-"'<* 



??.«»'?9!3^!!S>B!wr.f|^^ 



LMSC-A604141 



'^ee* "*" ; '^-.:-!»-'HifT.i^r«^>«*%— T^f-t^iir-^i. 



'i 






i'S::' 









f- 



:- « -, 



^ f 



Hi, 









J •ts 



S 6: -4* ^ii lS;s4* 



ii 






fc2 ^^ L' . Y, 









~C"~*- ■ - ~- 



i V 



. I A ;i 1 



CZ 






~r''**'^'^s^'5^^"*™*r''^"*'**'''v''''?'-^''^" ^'"^"v^fj/''"*- '*',■ '• ^ — ^ ■ 1 1 ,■ ■ 



U 






L, 



\ 
1 



■K«»^ 



^ititffeftji>*iflaitibi;^.:ua^^ 



.^ayb^buiausLt. 



0) 



o 



;=) 

Ml 

■ H 

X. 

CO 

1—1 

>^ 



X 
H 

!h 

I 

00 



tuO 

•r-t 



8-5 



LOCKHEED MISSILES & SPACE COMPANY 



»»Sj;4Hiy>;is tiiv.j.jp ij ,»ii»jijiiy i w,Mi..ijat.»iM»'!»ww 



LMSC-A604141 



'';'mv^SAiiiiS'^ii£rs.CAlsx)imw'-¥^^ -r. 






Kv" i. 



c ! 



I ! 















ti"'B-i;' S'iE--' iv'r... i^/ 



a 




i;w--'°.*iit2ji 



^, 



_4ti„»^-,.-=-. 



LII 



l. ' 



!' -*V 








/*^ 




IM^^'"-- 










■,...,| 






u 
I 

•4-) 

P 

bO 

• H 

to 

:^ 

r-t 
!>^ 



X 

a 
H 

nJ 

to 

I 

00 



bO 



I 



I 



LOCKHEED MISSILES 8c SPACE COMPANY 



^y HJ*»*-y. i '< - . > ayii,:".". ,-^'ifff^-ifjp!'^' r» r^. 



LMSC-A604141 






Iv 



•:; rv 



pif' 



': > 






.•rfi'fi ,'"- >'^uj'. .„jiSfe>;i 



•Iff '::, 






■/ 



•^, 



b. 



>■ -» 

f 



'"%''-i 



I 



T 






r r ^ 



i; 

I: 



\ ' 



5 ■ 



ML..j:jLf« ■i^u^xi^^t^^'liji^iilh^i^ 



Fig. 8-6 Vacuum Bake-Out Oven (Start Tank) 



8-7 



LOCKHEED MISSILES & SPACE COf'iPANY 



-^i^^^'^7r^'f^i!??^j-y^f>ft ^f » . '' -^ ' ^ V :>*-jm''y' ^ f ^ 



LMSC-A604141 



3, Model 8Z47 engine turbine overspeed and engine -mounted 
electronic gate checkouts 

4. PPS propellant isolation valve (PIV) functional checkouts 

Pneumatic Portion of Cart — Components of each of six separate gas systems 
(Fig. 8-7) are a relief valve, regulator, shutoff valves, vent valve, isolation 
check valves, pressure gage, flex hoses, and the necessary interconnecting 
plumbing. All system outlets protect the engine with integrally mounted- 
in-line 10-micron absolute filters. Duplicity of systenns is required so as 
not to use the same pressure system for checkout of both fuel and oxidizer 
conaponents. The pressure ranges of the six systems are to 30 (two 
systems), to 400, to 800, and to ZOOO psig, respectively. Maximum 
allowed inlet-supply pressure is 3000 psig. 

Electrical Portion of Cart — The electrical portion of the cart (Fig. 8-8) 
consists of circuitry required to perform functional checkout of the electronic 
gate. Electrical connections are provided to measure the engine "tell-tale" 
and hence verify proper relay operation within the unit. 

The electrical portion of the cart also supplies controls and connections to 
perform turbine overspeed checkout. Contained within the cart is a naeans of 
loading and mom.entarily shorting the secondary windings (Fig. 8-8) of the 
two isolation transformers which supply turbine speed signals to telemetry 
and the Agena status display panel. 

Switches, logic circuitry, and a 28-vdc timer are also mounted in the cart and 
are utilized to time the opening and closing of the PIV's. A design requirement 
of the timing circuit is that neither equipment malfunction nor operator error 
shall cause damage to the propellant isolation valves. 



8-8 



""^^/f!^ ! i .t<i p w |i |i..yj|nu i' l,Wj| ji ^«tj,<jyj^y i; 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 




o 

10 

o 
O 

■»-> 



o 

D 

U 



c 
o 

• H 
CO 
t— I 

:3 
o 

u 



o 

CI 

o 

■(-> 

u 
o 



I 

00 

•i-i 



8-9 



I 



LOCKHEED MISSILES & SPACE COMPANY 



T*'^Ui.'t;'Wl'!!l'*fl»L! 



r-rw^Kf ^. 1 1 II I i Jtnj^,i jm i ii i ni ! *! -w p imj i j^y w^y i yw j ie wM i jJAijjw. -v- ■^»i^'v - ^^i» ji w yj^ *t .j/,, ! v * *>u» > r,- ' v'ii -^^ ^* f yjr r .' T i | tiw « f^ !j ?> w j^.^ ' V ' ^f?>^'T^^'T^'- r-? ' ^^, * t ' » v .« ' ^ «' * '^ ' ?^gy y-^^-t y? ^x^ 



LMSC-A604141 




w 
CO 

PI 
o 

• H 

CO 

t-i 

P. 
o 
u 

u 
S 

• H 
iH 

0^ 

V.I 

O 

•J o 
M CO 

o c 
(1. ° 

o S 

•rH O 

^ -^ 
M O 



5 ?J 



00 
I 

00 
bb 



8-10 



LOCKHEED MISSILES & SPACE COMPANY 



^ ' ?^ a P .J ?^ ' ~j i! j P *^J ?^rT«g?rag^ ^ T^--?^MTW?W« 



ffi-'T?'jt?* .r ■fWJVtT.''J ''?y '<f f j^ fr-^<;y- ■■■-■->r^y7*ty .. »y/ng!;«»?fr 



rwwni^^a 



LMSC-A604141 



The following tools are supplied in kit form with the PPS checkout console and 
are used during functional flush and static leak checks of the Model 8247 
engine: 

1- Thrust Chamber Throat Plug — for plugging the thrust-chamber 
throat section during thrust-chamber leak check 

2. Turbine Exhaust Duct Plug — for plugging the turbine exhaust-duct- 
exit section during GG and exhaust-duct leak checks 

3. Fuel Pump Inlet Cap — for plugging the fuel pum.p inlet during fuel 
system pressure and leak checks 

4. Oxidizer Pump Inlet Cap — for plugging the oxidizer pump inlet 
during oxidizer system pressure and leak checks 

5. Pilot-Operated Solenoid Valve (POSV) Checkout Tool - for 
determining poppet leakage of the POSV 

6. Oxidizer Valve Actuation Tool — for attaching to the inlet of the 
oxidizer valve when checking actuation pressure of the valve 

7. Oxidizer Venturi Flushing Adapter —for replacing the venturi 
during oxidizer GG system flush. 

8. Fuel Venturi Flushing Adapter — for replacing the venturi during 
fuel GG system flush 

9. Oxidizer Valve Flushing Adapter —for providing assistance in 
overcoming oxidizer valve spring force and providing a port for 
flush fluid flow 

10. Thrust Chamber Throat Protector — for protecting thrust-chamber 
throat lining during flush. 



8-11 



■ LOCKHEED MISSILES & SPACE COMPANY 

M ' J^- '| » l ^^ ' :J W!g'^Wl gU !! . '' ^r^W#'^^^^' ! WP !Wj^WBy 'MI,'J ' ■^ 



{'1 
LMSC-A604141 y 



8. 3 START TANK LOADING CARTS 

8. 3. 1 Oxidizer Start Tank Loading Cart 

:ir This ca.t consists of five system, (F.g. 8-9). T.ey are: 

, p,essHiStion.Sour«llste™. which consists of the valves and 

, Xra„sferT>ie.P££HHHi«£ESii2:i' "'''=*^ ^^ ""' *° '"Z' ! 

into the start tank ^-t <.^ +r^ 

3 Fluid Transfer System, which has valving and flex Unes . UU.ed to 
St7^;;rT;i:;;^;7;^ropellant mto and from the AGK-caUhrated 

trpsqel to the- start tank i • .11 

4. IJ^^^UkGa^BeUowsJ^stes^ which is used to supply •■staC.ng 

pressure to the gas side of the start tanlc bellows immediately 
prior to Start tank load 
5. Star^rankLiquidPu^^ which is used to purge the Uquxd 
-;r^;^[^^:Z7^^;:[^^^Z^^:^ start tank fill valve and out the 
liquid bleed port mounted on the start tank. 

^v.. r^;,rt have in-line 25-naicron absolute filters, and 
All liauid systems on the cart nave m imc u. 

.v.. cart have 10-micron absolute filters to provide max- 
all gas systems on the cart nave lu 

imum cleanliness to engine start tanks. 
8.3.2 Fuel Start Tank Loading Cart 

, , ^- ..r-i- IViP 8-3) is a portable unit (equipped with 
The fuel start tank loading cart (Fxg. » ^M^ P 
tires) which contains valves, regulators, relief valves, gages, filters. 



8-12 



LOCKHEED MISSILES a SPACE COMPANY 



I 






LMSC-A604141 



I 







> 
u 

CO 

H 

u 

4-1 

w 

in 
flJ 

• rH 
•H 

O 



I 

00 



bO 



8-13 



■ LOCKHEED MISSILES & SPACE COMPANY 

i .tjjj^. i y. f i. ii ||^yjll » )i.,<<njjM i jjt> t j l|;i y;,iJ|iig{ ii ^^ 



LMSC-A604141 



1. 



2. 



oxidizer storage vessel (volumetrically calibrated), calibrated gas storage 
vessels, and attendant interconnecting plunabing sufficient to accurately load 
the Model 8247 engine fuel start tank and accurately precharge both fuel and 
oxidizer start tank gas sides. This cart consists of six systems (Fig. 8-10). 
They are: 

Pres surization Source System, which consists of the valves and 
regulators necessary to reduce and regulate source pressure 
Transfer Tube Pressurization System , which is used to force, by 
15-psi regulated gas pressure, a measured volume of propellant 
into the start tank 

3. Fluid T ransfer System, which has valving and flex lines utilized 
to effect the transfer of propellant into and from the AGE -calibrated 
vessel to the start tank 

4. Start Tank Gas Bellows System, which is used to supply "stacking" 
pressure to the gas side of the start tank bellows immediately prior 
to start tank load 

5. Start Tank Liquid Purge System , which is used to purge the liquid 
side of the start tank through the start tank fill valve and out the 
liquid bleed port mounted on the start tank 

6. Fuel and Oxidizer Start Tank Precharge, which consists of two 
accurately calibrated gas volumes and is used to load a pre- 
determined mass of both the fuel and oxidizer start tanks. 



All liquid systems on the cart have in-line 25-micron absolute filters, and 
all gas systems on the cart have in-line 10-micron absolute filters to provide 
maximum cleanliness to engine start tanks. 



8-14 



LOCKHEED MISSILES & SPACE COMPANY 



...■"rfifi-rs'jp"*. r-?^*- 



LMSC-A604141 



a 



< 

o 







9 ^ 






iTAT 

AGE 

RATI 










n 




o m 


< 

o 




tn K _l 




liJ O < 






t- u. U 


to 






< 






o 


n 




^ 


r 


IjO 


z 
< 


(1 


X- 


l- 


LiJ 


. 



9 oj 



< 
> 







> 



4-> 

t— ( 



I 

00 



8-15 



LOCKHEED MISSILES & SPACE COMPANY 



W!Wj. g . ■■■ a, B ? ' \KK ^^-■■^ ■^ ^ •■ii w iw fi^^ii^^^ ' ^r . ^v. ^^^^ 



?^aH'iFi?^r*-*rrr''^'*T-'i 



PVr*:ff'g t:yt^ '»* ^ , j ? W 7 ' '^:;^^''^''* **'*s^5TT'''^'3i!^^ 



LMSC-A604141 



8. 4 AUXILIARY FLUSHING UNIT (START TANK FLUSHING) 

Fuel Start Tank Flushino ; — The fuel start tank flushing portion of the cart has 
one storage tank containing methyl alcohol which is pressure fed to the liquid 
side of the fuel start tank. This tank is cycled in a manner sinnilar to that 
of the oxidizer start tank —i.e., "pumping" the bellows until all traces of 
UDMH have disappeared. 

Oxidizer Start Tank Flushing — The oxidizer start tank flushing portion of 
the auxiliary flushing cart (Figs. 8-4 and 8-5) incorporates storage tanks 
containing heated inhibited water, methyl alcohol, and methylene chloride. 
Regulators, valves, and interconnecting flex lines are used to pressure feed 
the tri-flush cleaning agents into the liquid side of the oxidizer start tank. 
While these fluids are being individually fed into the tank, another AGE 
system supplies regulated nitrogen to the gas side of the start tank for 
alternately pressurizing and venting, thereby causing the bellows to move and 
"pump" the cleaning, agent from the tank inlet (liquid-fill valve) and out 
through the tank-mounted liquid-bleed port. The miore highly corrosive 
IRFNA requires three flushing nnediums to gurantee a clean system as 
opposed to the fuel systenn which requires one cleaning agent. 



8-16 



LOCKHEED MISSILES & SPACE COMPANY 



wi«!Eas£Neiw=*MrV5«rw'':^J*!9SF^'': 



LMSC-A604141 



8. 5 VACUUM START TANK BAKE-OUT OVEN 

When flushing of both start tanks has been completed, the tanks are removed 
from the flushing unit, and with all fittings and ports left open, the tanks are 
placed in a specially provided vacuum bake -out oven (Fig. 8-6). Then the 
tanks are baked for four hours at 28 in. of vacuum (approximately 0.7 psia). 
The baking accomplishes final drying of all start tank internal surfaces. 

8. 6 MAIN SYSTEM FLUSHING 

Main fuel and oxidizer system flushing is accomplished in a manner similar 
to Model 8096 engine flushing and will not be specifically discussed in this 
report. Any special hardware required to perform the Model 8247 engine 
flush accompanies the PPS checkout console and is listed in Paragraph 8.2 
of this section. 



8-17 



LOCKHEED MISSILES & SPACE COMPANY 

iM II I ■Mill ■■■■i^«n iiJBi.i jT"r"~T^' ""f n'""'' 



LMSC-A604141 



BIBLIOGRAPHY 

1. Agena D Propulsion System for NASA Programs Engineering Analysis 
Report , LMSC-A372108, 2 April 1963, Confidential 

2. Recommended Method for Determining Turbine Back-Pressure, 
LMSD-419540, 22 December 1958, Confidential 

3. Determination of Vacuum Thrust Coefficient, Bell Aerospace Corp. , 
8048-982-005, 27 October 1958, Confidential 

4. Sutton, G. P. , Rocket Propulsion Elements, Second Edition, Wiley and 
Sons, 1956 

5. Marks, L. S. Mechanical Engineer's Handbook , Fourth Edition, 
McGraw Hill, 1941 

6. Model Specification Engine Rocket, Liquid Propellant USAF Model 



XLR-81-BA-13, LMSC- 1414463, 22 July 1963, Confidential 

V- BAG Model 8247 Service and Maintenance Handbook Liquid Propellant 
R ocket Engine, Report No. 8247-954201, 31 January 1964 

8. Gemini Agena Target Vehicle Familiarization Handbook , LMSC-A602521, 
1 April 1964 

9. Gemini Agena Target Vehicle Parameters Handbook, LMSC-A374366, 
1 May 1963, Confidential 

10. Propulsion System Test Requirements for Manufacturing an d Final 
Acceptance Tests S-OIB, LMSC- 1414805, 31 October 1963 

1 1- NASA Propulsion Subsystem Test Requirements Specification (SSOIB) , 
LMSC-1415487, 27 April 1964 

12. Telemeter System. Instrumentation Schedule, Agena /Genaini, 
LMSC-1352265, 6 March 1964 

B-1 



■ LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



13. Satellite Pyrotechnics Course No. SS 154 , LMSC-A067093, 9 July 1963 

14. Retrorocket (Design Specification ), LMSC- 1414701A, 23 July 1961 

15. Bolt, Explosive Dual Cartridge (Design Specification) , LMSC- 141 2344D, 
16 June 1961 

16. Specification, Squib - Pressure - Altitude, LMSC- 141 5503, 1 November 
1962 

17. Specification - Self Destruct Charge, LMSC- 1067259H, 1 November 1961 

18. Specification - Umbilical Coupling, Quick Disconnect, Helium and 
Nitrogen Fill, LMSC- 106873 ID, 23 June I960 

19. Process Specification (LMSC), Lines and Fittings, Gas and Liquid; 
Cleaning of , LAC (S) 1481B, 8 September 1963 

20. Status Report, Gas Fill Valve , LMSC-A635670, 25 February 1964 

21. Requirement Specification, Agena Target Vehicle Propulsion System Test 
Requirements, LMSC- 141 6537, 17 April 1964 

22. Weight and Performance Status Report - Gemini - Agena D , SP-129-64-2, 
1 April 1964 

23. Nozzle Extension BAG Model 8247 Design Analysis Report, Report No. 
8247-910015, 4 October 1962 

24. Design Control Specification, Remote Operated Quick Disconnect Type 
Pr opellant Umbilical Connectors , LMSC- 1067287, 20 July 1964 

25. Specification - Propellant Vent Couplings , LMSC-1510458 E, 26 February 
I960 

26. Instrumentation Description Agena Target Vehicle , Technical Memorandum 
91-35-2, 13 December 1963 

2'''. Propulsion System, Secondary, Liquid Propellant, BAG Model 8250 

Engineering Analysis Report, Report No. 8250-910002, 4 January 1964 



B-2 



LOCKHEED MISSILES & SPACE COMPANY 



LMSC-A604141 



28. Vehicle Umbilical and Test Plug Pin Assignments, LMSC-A068386, 
3 April 1964 

29. Ag ena Target Vehicle Status Display P anel Description, 
LMSC-A602638 

30. Propulsion System Analysis and Preliminary Desig n Report for 
Agena/Gemini Mission, Report No. LMSC-A055780, 27 July 1962, 
Confidential 



I 
I 



B-3 



LOCKHEED MISSILES & SPACE COMPANY 

il il l |»j ! W i u8jt,l » WU ! i» i y^W.M| |ii |,MltW. i if)t"M»W ! 



.'■■■'»r'Tw^l«p ¥»' W' t>^V H " V'< f! l>y''i^'W'r- tiWWy,» ! ^W '' «»^ ^^